If you try to go on a trip in your automobile, you are not going to get very far if there are no gasoline stations to feed your auto. Or restaurants to feed you. Or auto repair shops. This is what is called infrastructure. In the same way, if you want a rocketpunk future, you are going to need some infrastructure in space or your spacecraft are not going to get very far either.
Having said that, creating these pieces of infrastructure will be very expensive. It will be difficult to fund them. And you can be sure that whoever manages to build them will have iron control over who is allowed to use said infrastructure. And how much will be charged as a fee to use it.
It is possible to use spacecraft without any of this infrastructure, but it will be much more difficult. The owners of the infrastructure will probably adjust the fees so it will be cheaper to use their services, but only barely. The macroeconomics of the solar system will be vital to making all of this work. Especially the various business opportunities.
And of course an entertaining series of future histories can be postulated using various initial conditions. Does one national government have a monopoly? Two or more governments? Does one privately owned corporation have a monopoly? Two or more corporations? Or a several governments and several corporations?
For most missions, almost half of the delta-V budget is used up in the first 160 kilometers or so, the lift-off from Terra's surface into Low Earth Orbit. This is the reason behind Heinlein's "halfway to anywhere" comment. In dollar terms, the Russian Proton would cost about $5000 per kilogram boosted into LEO while the Space Shuttle would cost about $18,000 per kilogram. Actually if you factored in all the shuttle's design and maintenance costs, the real price was closer to $60,000 per kilogram. NASA was hoping that the shuttle would cost more like $1,400/kg (in 2011 dollars), though part of the over-run was due to the multi-year interruptions in launches following Shuttle failures.
See the section on Laser Launching.
As the delta-V for a mission goes up, the amount of propellant required goes up exponentially (or looking at it another way: the amount of payload shrinks exponentially). Large amounts of propellant are expensive, but the higher the mass-ratio the higher the likelihood that the spacecraft will not be resuable. Propellant expense is bad enough, but that's nothing compared to having to build a new spacecraft for each mission. Increasing the mass ratio means things like making the walls of the propellant tanks thin like foil, and shaving down the support members so they are fragile like soda straws. With such flimsy construction it does not take much normal wear-and-tear to turn the spacecraft into junk.
Having a propellant depot at the mid-point of a round-trip mission cuts the required delta-V in half. Instead of the spacecraft having to lug enough propellant to go to Mars then return to Terra, it only carries enough for half the trip but re-fuels (re-propellants, or re-remasses) at the halfway point. And when you are dealing with exponential growth, cutting the delta-V in half cuts the propellant amount much more than half.
Indeed, Rick Robinson noticed that with access to an orbital propellant depot, most cis-Lunar and Mars missions are well within the delta-V capabilities of a sluggish chemical rocket engine. You do not have to use a nuclear thermal rocket. Hop David noticed this as well. Dr. Takuto Ishimatsu's ISRU optimization algorithm calculated that NASA's Mars Reference Mission was more optimal with no NTR but with ISRU ("optimal" defined as "requiring less mass boosted from Terra into LEO").
This is also an argument for orbital propellant depots in Low Earth Orbit. Remember that once the rocket has traveled from Terra's surface into LEO, you are "halfway to anywhere". This means for a one-way trip, LEO is the mid-point of the mission.
Now to make this work, in addition to the depots you will need sources of propellant and tankers and lighter to keep the depots filled up. Water and hydrogen propellant is available in such places as the lunar poles, asteroids, and perhaps the Martian moons Phobos and Deimos. Dr. Takuto Ishimatsu developed an algorithm to optimize placement and supply of ISRU orbital depots.
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.
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 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
Demand Assumptions and Methodology
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
- Follow same orbital distribution
- Lifetime of 20 years
- 10X current maneuverability
- 10% share of all ISRU water
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
Supply Chain Road Map
Calculating Supply 1: The Moon
Supply Chain Road Map
Lunar Assumptions and Methods
- 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
- 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
- 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
Calculating Supply 2: Near Earth Asteroids
Supply Chain Road Map
Near Term NEA Assumptions and Methods
- 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
- 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
Near Term NEA Model
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
Long Term NEA Assumptions and Methods
- 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
- 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
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
Calculating Supply 3: The Asteroid Belt
Supply Chain Road Map
Asteroid Belt Assumptions and Methods
- 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
- 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
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
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
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.
Kuck Mosquitoes were invented by David Kuck. They are robot mining/tanker vehicles designed to mine water propellant from icy dormant comets or D-type asteroids and deliver it to an orbital propellant depot.
Form follows function. So it is unsurprising that the Kuck Mosquito resembles an Enterobacteria phage T4 virus. Only difference is that the virus is injecting, while the Mosquito is sucking out.
Kuck Mosquito Propulsion H2-O2 Chemical Specific Impulse 450 s Exhaust Velocity 4,400 m/s Wet Mass 350,000 kg Dry Mass 100,000 kg Mass Ratio 3.5 ΔV 5,600 m/s Mass Flow 49 kg/s Thrust 220,000 newtons Initial Acceleration 0.06 g Payload 100,000 kg Length 12.4 m Diameter 12.4 m
Deimos, the outer moon of Mars, is possibly the most accessible source of water to LEO. Lewis has shown the delta-V to go from LEO to Deimos is less than that needed to land on Earth's Moon. Partial loss of velocity at Mars might be obtained by a shallow dip into the Martian atmosphere. The delta-V to return from Deimos to HEEO (Highly eccentric Earth orbit) is very small. The travel time is roughly two years. The Moon may be used as an aid to accelerate and decelerate a vehicle as it leaves LEO and arrives at HEEO. Shallow penetration of the Earth's atmosphere may be used to loose velocity and aid in capture into HEEO.
Outbound Inbound Body delta-V Surface to LEO (m/sec) time of flight (d) delta-V LEO to Surface (m/sec) time of flight (d) Phobos/Deimos 5600 270 1800 270 Moon 6000 3 3100 3 Mars 4800 270 5700 270
(ed note: the important part is LEO to Deimos Surface is deltaV=1800 m/s and 270 days transit, Deimos Surface to LEO is deltaV=5600 m/s and 270 days transit.)
A disadvantage of Deimos is the 26 month delay between launch opportunities.
Fanale calculates that ice should exist at a depth of 100 meters at the equator and at a depth of 20 meters at the poles of Deimos. Thus, the drilling equipment proposed in 1995 by Kuck should be able to reach ice at or near the poles, but not near the equator.
To move 100 tonnes of water ice from Deimos to LEO will require 250 tonnes of water ice for propellant (Z). Thus, in order to leave Deimos 350 tonnes must be propelled from the surface. A 1,000 cubic meter collection bag should be large enough to contain the 350 tonnes of ice, cuttings & other precipitates.
(ed note: 100 metric tons payload + 250 metric tons propellant implies mass ratio of about 3.5, since engine and structural mass are a small fraction of this (e.g, by the table below, the drilling equipment is 0.3 metric tons). If it electrolyzes the propellant into O2 and H2 and burns it as chemical fuel with a specific impulse of 450 seconds, this would give a delta-V of around 5,500 m/s or so.)
Table 1. Mass of Drill and equipment for the Deimos version of the drill presented in "Exploitation of Space Oases" presented at Princeton May 1995. The total mass is in grams. The drill pipe is titanium for lightness and chemical resistance to corrosion. Down The Hole Hammer Drill, Titanium drill pipe & accessories L (mm) OD (mm) ID (mm) Weight (grams) Number Weight (grams) Ti Hammer DTH 210 16 233 3 699 No Under-reamer Guide 78 20 49.4 3 148.2 Yes 117 30 67.1 3 201.3 Yes Under-reamer 15 27 20 10 200 No 20 37 36.5 10 365 No Casing Shoe 21 24 16 10 160 No 26 35 29 10 290 No Tubing 2000 16 14 425 325 138125 Yes Casing 2000 22 20 595 100 59500 Yes 2000 32 30 1299 60 77940 Yes Collar Pipes 1000 43 40 1374 10 13740 Yes Total 291368.5
The images below are details of the "Spider" water harvester carried by the Robot Asteroid Prospector. It performs much like the business end of the Kuck Mosquito.
Takuto Ishimatsu's Ph.D. thesis is titled Generalized multi-commodity network flows : case studies in space logistics and complex infrastructure systems (abstract here). Basically: manually chosing where to place in-situ resource allocation depots so as to get the maximum benefit is a very very hard problem. Wouldn't it be nice to create a computer program that could automatically find the optimum solution?
This is very relevant to our interests.
For the Apollo missions NASA used a "carry-along" strategy, where all vehicles and resources traveled with the crew at all times. Along with the horrific propellant cost to boost all of this from Terra into LEO. For the International Space Station NASA adopted a "resupply" strategy. This also has horrific boost cost plus it requires a close resupply source (Terra).
The resupply strategy ain't a gonna work for a Mars mission (as Terra gets further and further away), so the conventional view was to use a carry-along strategy. Dr. Ishimatsu examined NASA's Mars Design Reference Architecture 5.0 (plus addendum 1 and 2).
As you know from reading this section the way to avoid the horrific boost costs is in-situ resource utilization: travel light and live off the land. The problem is figuring out what is the best placement of in-situ mining, refining, and orbital depot assets.
Dr. Ishimatsu's software determined that using lunar propellant mines and tankers would cut the cost of the conventional NASA Mars mission by a whopping 68 percent!
It is very similar to the military. The old bromide is that amateurs talk about battle tactics while professionals talk about logistics. Well, deep space exploration is going to require a well-planned logistics strategy as well.
Dr. Ishimatsu examined several prior solutions, but all either were not scaleable as the mission complexity increased, required the user to pre-define the logistics network (i.e., solve the problem manually), or were not capable of doing optimization with no human input.
Dr. Ishimatsu used Dale Arney and Alan Wilhite's technique of modeling space system architectures using graph theory. The nodes are physical locations in space wihle the arcs (connections between the nodes) are possible movements or transports between nodes. Note that arcs are one-way, an arc going from node A to node B is totally different from an arc going from node B to node A. This is because one can, for instance, use aerobraking to traverse an arc going from LEO to Terra's Surface, but one cannot use aerobraking to go from Terra's surface into LEO.
To allow for the optimal solution, it is best to include as many nodes and arcs as possible. The optimizer obviously cannot use arcs and nodes that are not present. If the optimal solution requires use of a missing node or arc, it will not be found.
One peculiarity is that you use an arc that starts and ends on the same node to model a node that is a resource processing facility. This is required in order to allow the optimizing mathematics to work. These are called a "graph-loop", "self-loop" or a "buckle".
Another peculiarity is having several arcs between a given pair of nodes. For instance, if the mission could move items between node A and node B by either chemical rockets or nuclear thermal rockes, each rocket type would have its own arc between node A and node B. This is because the two rocket types have different specific impulse and thus different propellant consumption. Additional arcs will be required for the same rocket type if it has different delta V usage choices. For instance, a nuclear thermal rocket can do either an economical burn with a long time of flight or an expensive burn were more propellant was expended in order to reduce the time of flight. There will also be an additional arc where aerocapture is possible.
Basically the multiple arcs allow the optimization to explore multiple mission choices. One choice per arc.
These multiple arcs between a given pair of nodes are called "parallel arcs."
For logistics calculation, you state the mission as a set of demands at certain nodes in the network. A demand for "plantISRU" at the LSP node corresponds to a lunar mission to transport an in-situ resource allocation industrial plant to the lunar south pole.
Dr. Arney modeled the propellant required in a mission as costs on a given arc. But Dr. Ishimatsu found it more useful to model the propellant required for all subsequent stages of the mission as payload on a given arc. In addition, since in-situ resource allocation (ISRU) allowed propellant and other resources to be generated at other nodes besides Terra, it made sense to model propellant as commodities included in the flow variables rather than as costs like Dr. Arney did. This allows formulating the problem as a multi-commodity network flow, with some commodities coming from Terra and others from ISRU sites.
The optimization problem becomes finding the best routes in the network that satisfies the mission demands while also meeting certain constraints (i.e., figuring out which nodes and arcs to use). The result will tell you "where to deploy what."
The program is trying to optimize TLMLEO, which is Total Launch Mass from Terra to Low Earth Orbit (LEO) required to set up the entire logistics network. The program is trying to find the solution with the lowest TLMLEO.
Note that there are lots of other things that could be optimized for, but this system only optimizes TLMLEO. Other things that might be optimized include:
- Development, Test, and Evaluation cost of the various components (ISRU and orbital propellant depots will require lots of expensive R&D)
- Number of rendezvous and refueling events (the more, the higher the chance of a malfunction or accident)
- Complexity (the more complicated, the more potential points of failure)
For the nitty-gritty mathematical details of the optimization, please refer to Dr. Ishimatsu's thesis. It contains lots of calculus and matric algebra which makes my head hurt. There are matrix multiplications for flow equilibrium, flow transformation, and flow concurrency.
As a case study, Dr. Ishimatsu ran NASA's Mars Design Reference Architecture 5.0 through his software.
In the model, everything that travels from node to node is a "commodity", even the crew. The 20 commodities are listed in the table below. Each commodity has a flow and demand all measured in kilograms.
- crew (traveling to Mars)
- crewRe (returning to Terra)
Commodity "crew" represents the crew traveling from Terra to Mars while "crewRe" represents the crew returning to Terra. A self loop on Mars transforms crew into crewRe, enforcing the rule that the mission is a round trip. This is a mathematical trick that allows the optimization math to work.
The Resources catagory includes the rocket propellants, crew provisions, and crew wastes.
The Infrastructure catagory includes habitation facilities, ISRU industrial plants, and ISRU spares.
The Transporation catagory includes vehicles, propulsive elements, and non-propulsive elements. "InertX" means "rocket engine utilizing propellant X" while "TankX" means "tank full of propellant X. The three engines are: chemical liquid oxygen + liquid hydrogen, chemical liquid oxygen + liquid methane, and nuclear thermal rocket. Note that NTR can use any of the three tanks as propellant, the others require tanks of each of their named propellants. For NASA reasons, the NTR is not allowed for lift-off or landing on a planet, and aerocapture is allowed for unmanned cargo missions but not allowed for manned missions.
Solar electric rockets were not included because they require a different way of defining the arc parameters.
For the Mars mission it requires a demand for "habitat" at GC and a demand for "crewRE" at PAC. This translates into a mission to send a crew of six and a surface habitat to Mars Gale Crater, the crew becomes crewRE (crew ready to return to Terra) on Mars after a 540 day stay, which forces a mission to send the crewRE from Mars to Terra Pacific Ocean Splashdown.
Dr. Ishimatsu used the graph below, which does show the self-loops but only shows a single arc even when parallel arcs are present. Otherwise the diagram would be an unreadable mess. The full graph has 16 nodes and 598 arcs. There are self-loops at LSP (Lunar south pole), DEIM (Deimos), PHOB (Phobos), and GC (Mars Gale Crater).
ISRU availability/technology have the folloiwng assumptions:
- Lunar ISRU can produce O2 from regolith or H2O from water ice at a rate of 10 kilograms per year per unit plant mass while requiring spares of 10% of plant mass per year.
- Mars ISRU can acquire CO2 from the atmosphere or H2O from water ice with the same production rate and spares requirement as those for lunar ISRU.
- Mars CO2 can be converted into CH4 and H2O via the Sabatier reaction or can be converted into O2 via solid oxide electrolysis.
- Electrolysis of H2O and pyrolysis of CH4 are assumed to be available along with lunar/Mars ISRU
All these chemical reactions are modeled as an optional self-loop.
First, a "baseline" problem is defined and sent through the program for a solution. This is a simplified problem whose solution will be used to measure the results of altering various parameters. Among other things the baseline problem has the propulsion system modeling simplified. For the details about the baseline problem, please refer to Dr. Ishimatsu's thesis
Dr. Ishimatsu rubs salt in the wound by cheerfully telling us "Using MATLAB 8.3 (R2014a) with CPLEX 12.6 on an Intel R CoreTM i7-2640M CPU at 2.80 GHz, one run of the optimization model takes approximately 12 seconds for preprocessing and 1.2 seconds for optimization (TLMLEO minimization)." He did a test validation by constraining the model to NASA's Mars Reference Mission, the results were practically identical.
The baseline solution has a TLMLEO of only 271.8 metric tons, a 68% savings from the NASA Mars Reference Mission NTR scenario, and a 78.3% savings from NASA's chemical/aerocapture scenario.
Then the user can alter various propulsion parameters and measure the results against the baseline solution. The other parameters and assumptions remain the same. The surprise here is that NASA's Mars Reference Mission's reliance on nuclear thermal rockets is sub-optimal. LOX/LH2 chemical engines are superior, if you include ISRU (which NASA did not). The massive amounts of oxygen and hydrogen produced by the Lunar ISRU more than makes up for the relatively low specific impulse of the chemical rocket.
The arc from Kennedy Space Center (KSC) to Low Earth Orbit (LEO) has by far highest delta V cost: 9.8 km/s. This is the mathematical way to model Heinlein's "Halfway to Anywhere" observation. The emergent property produced by optimization is the need for in-situ resource utilization.
Now the user can alter various ISRU availability scenarios and measure the results against the baseline solution. The other parameters and assumptions remain the same.
Spacecraft will need maintenance, and some will occasionally need major repairs due to damage (or gunfire). Obviously repairs will be eaiser if the engineers can perform them while wearing shirt-sleeve clothing instead of encumbering space suits. Most spacesuits raise the energy expenditure to do a task by about 400%.
Surrounding a spacecraft with an atmosphere can be easily done if:
- the spacecraft is near a planet with a ground repair dock
- the planet with the dock also has a breathable atmosphere
- the spacecraft is designed to land on a planet with an atmosphere, that is, the ship is not an orbit-to-orbit type or can only land on airless planets
- the damage to the spacecraft is mild enough that it is capable of landing
If any of these are not true, the ship will need an orbital drydock.
This is a space structure big enough to hold the spacecraft, capable of pressurizing the interior to shirt-sleeve conditions, and full of repair-crew and their tools. Probably inside or near a space station.
Locations too impoverished to afford such structures will just have to make do with space suited crews or remote drones with waldoes. Such facilities are called orbital wetdocks.
The non-offical term for both is "spacedock."
James Snead has written a few paper about space infrastructure. Most interesting is Architecting Rapid Growth in Space Logistics Capabilities. On page 23 he gives an example of an orbiting space logistics base, including a space dock. Refer to that document for larger versions of the images below.
...the space logistics base’s functions are: (1) housing for travelers and operating crews; (2) emergency care; (3) in-space assembly, maintenance, and repair; and (4) materiel handling and storage.
The example space logistics base consists of four elements. At the top in Fig. 10 is the mission module providing the primary base control facility, emergency medical support, and crew and visitor quarters. The personnel quarters are located inside core propellant tanks that are retained from the SHS used to launch the mission module. The overall length of the mission module and propellant tanks is approximately 76 m (250 ft). Solar arrays and waste heat radiators (shown cut-away in Fig. 10) are mounted on a framework surrounding the mission module to provide additional radiation and micrometeoroid protection.
The second element consists of twin space hangars. These serve as airlocks for receiving spaceplanes and provide a pressurized work bay for conducting on-orbit maintenance of satellites and space platforms.
As shown in Fig. 11, the space hangar consists of a structural cylindrical shell 10 m (33 ft) in diameter, a forward pressure bulkhead containing the primary pressure doors, and an aft spherical work bay. These elements, which define the primary structure, would be manufactured as a single unit and launched as the payload of an SHS. The large, nonpressurized, space debris protection doors would be temporarily mounted inside the hangar for launch and then demounted and installed during the final assembly of the hangar at the LEO construction site. All of the other hangar components would be sized for transport to orbit in the cargo module of the RLVs and then taken through the hangar’s primary pressure doors for installation.
Future logistics supportability is a key feature of this hangar design. The size, weight, location, and access of the internal hangar components enables them to be inspected, repaired, and replaced without affecting the primary structural / pressure integrity of the hangar. With the exception of the space debris protection doors, this would be done inside the hangar when it is pressurized. The ISS-type airlock and space debris protection doors, although mounted externally, would be demounted and brought into the hangar for inspection, maintenance, and repair. For the repair of the primary pressure doors, they would be demounted and taken into the spherical work bay or the other hangar for servicing.
The hangar’s design enables both pressurized and unpressurized hangar operations to be undertaken simultaneously. When the main hangar deck is depressurized to receive cargo or spaceplanes, for example, pressurized maintenance operations would continue inside the 9.8 m (32 ft) diameter spherical work bay and the 2.8 m (9 ft ) diameter x 4.3 m (14 ft ) work compartments arranged along the top of the hangar.
Hangar operations in support of the passenger spaceplanes, as shown in Fig. 12, highlight the improvement in on-orbit logistics support enabled by the large hangars. After entry into and repressurization of the hangar, the passengers would disembark from the spaceplane. Support technicians, working in the hangar’s shirtsleeve environment, would inspect the spaceplane and, in particular, the thermal protection system for any damage to ensure that it is ready for its return to the Earth. While at the space base, the spaceplane would remain in the hangar to protect it from micrometeoroid or space debris damage. Minor repairs to the spaceplane could also be undertaken to ensure flight safety.
The third element is the air storage system. The prominent parts of this system are the large air storage tanks that are the reused core propellant tanks from the two SHS used to launch the twin space hangars. Besides storing air from the hangars, this system also: manages the oxygen, carbon dioxide, and moisture levels; removes toxic gases, vapors, and particulates; and, controls the temperature and circulation of the air within the hangar and its compartments.
The fourth and final element is the space dock. It would be constructed from structural truss segments assembled within the space hangars using components transported to orbit in the RLVs. The space dock would provide the ability to assembly and support large space logistics facilities, such as the space hotels and large manned spacecraft described in the following. It could also used to store materiel and as a mount for additional solar arrays.
The space hangars and space dock would enable traditional logistics operations of maintenance, assembly, and resupply to be routinely conducted in Earth orbit. This is an enabling capability necessary to become spacefaring and achieve mastery of operations in space.
The space logistics base would have approximately 20 personnel assigned. The tour of duty would be 90 days with half of the crew rotating every 45 days. Crew rotation and base resupply would require approximately 32 RLV missions per year per base with 8 spaceplane missions and 24 cargo missions. This would provide approximately 12,000 kg (26,000 lb) of expendables and spares per person per year. At $37M per mission, a ROM estimate of the annual transportation support cost per base would be approximately $1.2B.
While the LEO space logistics base would have sufficient housing capacity to support the 20 assigned personnel and a modest number of transient visitors, it would not be a primary housing facility. Since people cannot simply pitch a tent and “camp out” in space, establishing early permanent housing facilities is an important and enabling element of opening the space frontier to expanded human operations. The architecture of the Shuttle-derived heavy spacelifter and the LEO space logistics base was selected so that the first large space housing complexes, referred to as space hotels, could be constructed using the same space logistics base modules.
A composite illustration of the design, assembly, and deployment of the example space hotel is shown in Fig. 13. This hotel design is configured as a hub and spoke design with a long central hub and opposing sets of spokes attached to the central hub module. This configuration makes it possible to use variants of the space base’s mission modules and space hangars as the primary elements of the space hotel’s design.
Element 1, in Fig. 13, shows the start of the hotel assembly sequence. The central hub module, shown with the SHS’s core propellant tanks still attached, is being positioned at the space logistics base’s space dock. The central hub module would be a version of the mission module used in the space logistics base. Its design would include 12 docking ports around its circumference for attaching the spokes.
Element 2 shows the completed hub and one attached spoke. Two space hangars are located at the ends of the hub and the first spoke is shown attached to the central hub module. In assembling the hub, the core propellant tanks from the two SHS missions used to launch the hangars would be incorporated into the hub to provide additional pressurized volume. This approach would be also used for the spokes. Each spoke would consist of a generalpurpose mission module with the SHS’s core propellant tanks reused for additional pressurized volume. As with the mission module on the space logistics base, the spokes would be surrounded by solar arrays and waste heat radiators. This is what provides their “boxy” appearance.
Element 3 shows the completed 100-person space hotel with two pairs of spokes on opposing sides of the hub. This is the baseline space hotel configuration. Seven SHS missions would be required to launch the hub and spoke modules for the baseline hotel. One additional SHS cargo mission would be used for the solar arrays and waste heat radiators.
This design enables the hotel to be expanded to 6, 8, 10, or 12 spokes. Each spoke would require one additional SHS mission. The 12-spoke configuration would accommodate up to approximately 300 people. Each additional spoke would be tailored to provide a specific capability, such as research and development facilities, tourist quarters, office space, retail space, etc.
Element 4 shows the completed space hotel after being released from the space dock. It also shows how the hotel would rotate about the long axis of the hub to produce modest levels of artificial gravity in the spokes. At about two revolutions per minute, a Mars gravity level is achieved at the ends of the spokes. This use of artificial gravity enables the spokes to be organized into floors (Element 5 in Fig. 13). Each spoke would contain 18 floors with 14 of these available for general use and the remaining 4 floors used for storage and equipment. The spokes would be 8.4 m (27.5 ft) in diameter. This would provide a useful floor area of approximately 42 m2 (450 ft2) per floor. The total available floor area in the baseline configuration would be 2,340 m2 (25,200 ft2). The 12-spoke configuration, having 192 floors total, would have 3 times this floor area—7,026 m2 (75,600 ft2) or about 23 m2 (250 ft2) per person.
An estimate can be made of the number of guests visiting the hotel each year. Assuming a 3:1 ratio of guests to staff, approximately 76 guests would be staying each night in the baseline configuration and 228 guests in the full configuration. With one third of the useful floors configured as guest cabins, two cabins to a floor, each cabin would have a useful area of approximately 21 m2 (225 ft2).* With an average stay of one week, approximately 4,000 guests and 12,000 guests would visit the 4- and 12-spoke hotels each year, respectively.
If each passenger spaceplane carries 10 guests, approximately 400 and 1,200 RLV flights would be required each year. With an additional 25% required for staff transport and resupply, the 4-spoke hotel would require about 10 flights per week and the 12-spoke hotel would require about 30 flights per week. If the RLVs could achieve a one-week turnaround time, and allowing for one in five RLVs being in depot for maintenance, 12 RLVs would be required to support the 4-spoke hotel and 36 RLVs for the 12-spoke hotel.†
At the $37M per flight cost discussed previously for first generation RLVs, the per passenger transportation cost would be approximately $3.7M. With this transportation cost structure, a sustainable space tourism or space business market may not be possible. However, if a second generation RLV could reduce this cost by a factor of 10 to $0.37M per passenger, as an example, then an initial market demand for the baseline hotel may develop and be sustainable. In such case, the annual transportation revenue for the baseline hotel would be $3.7M x 500 = $1.9B and the 12-spoke hotel would be $5.6B.‡ This improvement in transportation costs would also yield a savings of 90%—approximately $1B per year—in the transportation costs to support the LEO space logistics bases. Human space exploration missions would also realize a significant cost reduction.
While developing a conceptual design of a space hotel would appear premature at this early stage of considering the architecture of an initial space logistics infrastructure, several important conclusions emerge that indicate otherwise:
1) Careful selection of the initial space logistics architecture can also establish the industrial capability to build the first space hotels necessary to enable the expansion of human enterprises in space.
2) A commercially successful space hotel will require second generation RLVs to lower further the cost of transportation to orbit.
3) In order for these second generation RLVs to be ready when the first space hotel is completed, the technology research investment would need to begin concurrently with the start of the detailed design of the initial space logistics systems. Conversely, for private investment to seriously consider building the first hotels, significant science and technology progress in developing the second generation RLVs must be demonstrated by the time the initial hotel construction contracts are made.
4) The benefits of reduced space transportation costs will also substantially lower the cost of operation of the initial elements of the space logistics infrastructure, leading to a likely increase in demand for more in-space logistics services.
5) Space hotels and second-generation RLVs may become an important new aerospace product for the American aerospace industry, establishing American leadership in this new and growing field of human astronautical technologies.
6) It is not unrealistic to expect, with the building of an integrated space logistics infrastructure, that hundreds of people could be living and working in space by 2020, growing to thousands of people by 2040 with many of these living in the first permanent orbiting space settlements.
* A standard cabin on the new Queen Mary 2 cruise ship has an area of 18 m2 (194 ft2). A premium cabin has an area of 23 m2 (248 ft2).
† Launch sites for these RLVs would be distributed around the world. This would allow operations at the space hotel to run 24 hours per day since there is no day and night in LEO.
‡ This further reduction could come about through the introduction of a spiral version of the first-generation RLVs where improvements to the high maintenance cost subsystems, e.g., engines, could substantially reduce the recurring costs. Another approach would be development of entirely new RLV configurations—perhaps a single-stage configuration—that would also result in a substantial reduction in recurring costs per passenger through subsystem design improvements and the ability to carry more passengers per trip. A key issue in both approaches is the amortization of the development and production costs. High flight rates, probably dependent on space tourism, would be required to yield an overall transportation cost sufficiently low to enable profitable commercial operations.
Space-based solar power (aka "Powersat") is one of those concepts that make one think about idealistic hippy futurists in the 1970's drunk on the idea of MacGuffinite that is also ecological and green. It is solar energy on steroids. By placing the solar collectors in orbit you get all the solar energy since ground based solar collectors can only gather the frequencies that our atmosphere is transparent to, and are hampered by rain clouds and/or the fact that it is nighttime.
You can get almost unlimited amounts of green energy: no nasty coal, oil, natural gas, or uranium is required. Groovy, man!
The fact that none of these exist today tells you that the difficulties are overwhelming.
At Terra's orbital distance from Sol the solar power flux is abotu 1,366 watts per square meter. Due to the inverse square law the power increases the closer you get to Sol (see table). And vice versa as well.
But we do not care about such stations, since other than being a species of MacGuffinite, it has nothing to do with spacecraft, right?
Au contraire! Read on to see how solar power stations can be a boon to spacecraft.
Let me take a minute to talk about solar moth rockets.
Remember the fundamental rule of rocket design: Every Gram Counts. The motivation behind the solar moth is "just imagine how much mass we could save if we eliminated the rocket engine from the design! Using the "magnifying glass incinerating an ant" principle, the solar moth utilizes a large mirror to focus the heat from the sun on the propellant, energizing it so it rushes out the exhaust bell, resulting in thrust.
It is a pity that solar energy is so diffuse around Terra's orbit. To really get worth-while amounts of heat, the solar moth will need huge mirrors. Which sort of eliminates the mass advantage of removing the engine.
That's where the powersat comes in. Have a powersat send power in a beam of microwaves and give the solar moth a microwave rectenna to receive the electrical energy! You will be using Beam-powered propulsion. The electricity can be used to heat the propellant. Suddenly your pathetically weak solar moth will be a super-powered muscle machine.
The joke is saying that electromagnetic and electrostatic propulsion systems (e.g., ion drives, VASIMR) are power hogs. Solar power arrays will have to be huge due to the low power concentration in sunlight (low as compared to the propulsion power demands). Nuclear reactors can easily supply the power but have ugly mass penalties. But the joke is on the cracking wag, beamed power is pretty much the same as an extension cord long enough.
Microwaves are difficult to focus, and the conversion from electricity to thermal energy has unavoidable inefficiencies. It would be nice to beam thermal energy instead of electrical energy. Can do: replace the microwave with a laser! Now you can use the same lightweight mirror on a solar moth, but with the much more intense radiant energy of a laser. It will be a laser thermal rocket. You can also use it on a solar sail craft and make it into a high-powered laser or photon sail. The advantage is that your delta-V capacity will be incredibly large. The disadvantage is that you are at the mercy of whoever owns the powersat.
If your laser thermal rocket is renting laser time from Beams-Я-Us, you better make sure that your bill is paid up. Otherwise they will pull the plug and your rocket will suddenly be powerless, and on a one-way ticket to nowhere. You might be able to limp along using solar power instead of the laser from Beams-Я-Us, but I would not bet your life on it.
And make sure you stick closely to the flight plan you filed with Beams-Я-Us, or they might have a problem keeping the beam aimed at you.
Beams-Я-Us might purchase their own laser thermal or laser sail ships. They will then be like a rail-road company, owning both the trains and the rails they run on.
As a matter of fact, the solar collector on the powersat will be much more effective if it was closer to the sun than Terra's orbit. Say: the orbit of Mercury. Now we're cooking!
And if we start beaming power to interstellar spacecraft, a stray beam might give some alien civilization a Wow! signal.
About this point all but the hopelessly dull are thinking "wait just a darned minute, what are the military applications?" Pretty good, actually. Have you ever heard of a Laser Combat Mirror? The laser-propulsion mirror eliminates most of the mass of the engine, the laser combat mirror eliminates most of the mass of a laser cannon. This will free up payload mass in the space warship so it can carry more of other kinds of weapons.
As the range increases the powersat beam rapidly becomes too diffuse to do damage due to diffraction. But a warship sporting a laser combat mirror can focus the seemingly harmless diffuse beam into an eye-searing ship-destroying pin-point. Again much in the same way that sunlight is too diffuse to harm ants, unless a naughty boy uses a magnifying glass to focus it into an ant-destroying death ray.
Powersat's weapon potential is so effective that they will probably be nationalized, removed from civilian hands, and turned over to the military.
And even without a fleet of warships with laser combat mirrors, a powersat all alone is a pretty fair orbital laser weapon. Without laser combat mirrors the range is limited, but within that range, whoo boy can they vaporize the heck out of enemy spacecraft, space assets, and even torch ground targets. Their huge solar panels make them fragile, but they can do plenty of damage before they are neutralized.
Not quite green ecological hippy anymore, is it?
G. Harry Stine's (writing as Lee Correy) wrote a rocketpunk novel called Manna. In the novel, the military branches of the space-faring nations would like to put five gigawatt High Energy Laser (HEL) satellites in orbit. Using fancy techniques they are powerful enough to get their weapon laser beam through Terra's atmosphere and incinerate targets on the ground.
The trouble is the militaries want the HEL beamer satellites to be stealthy. The root of the trouble is that a five gigawatt HEL beamer containing a +five gigawatt power source is about as stealthy as a New York 4th of July fireworks display.
If only the power source could be at some distance from the HEL beamer, sending the energy by electromagnetic waves. You know, the same way a powersat sends microwave energy to ground power stations... hmmmmmmm.
That would work, the HEL beamers could be stealthy little dastards with no nuclear power plant, but rapidly unfurling a powersat reception antenna when it came time to zap something.
Now comes a bigger problem. Nobody can build any powerstats.
Why? Well, no corporation is going to embark upon a multi-billion dollar project like a powersat without insurance. And no insurance company is going to underwrite a multi-billion dollar installation which becomes a military target the instant it redirects its power beam from a power station in order to energize a HEL beamer. Especially a military target so huge, easy to hit, and incredibly fragile as a powersat.
How to solve the problem? Well, since it is an insurance problem, there should be an insurance solution.
Through a series of international agreements, the Resident Inspection Organization (RIO) was formed. This international group regularly inspected all powersats, and insured that they stayed pointed at ground power stations. In exchange, the insurance companies would underwrite the powerstats. If any powersat started to energize something that might be a stealthed HEL beamer, RIO would sound the alarm to all the astromilitaries, presumable giving the military units enough time to blow the living snot out of the powersat.
Naturally the astromilitary of Nation Alfa would be angry at RIO squealing when astromilitary Alfa tried to energize one of their HEL beamers. But astromilitary Alfa would be vary grateful if RIO squealed about astromilitary Bravo, Charlie, Delta or Echo doing the same thing.
A good low-mass way to prevent cables from failing catastrophically is to use Hoytethers (cables that are elongated Hoytubes). Strengthening a cable by increasing its diameter quickly becomes too expensive in terms of mass. A Hoytether on the other hand is a low mass network of redundant cables that fails gracefully.
Momentum exchange Hoytethers were featured in the novel Saturn Rukh by Robert L. Forward.
If you were shipping asteroid ore (actually "ore" is not quite the right word but there isn't a good one) from Ceres to Terra (or manufactured goods going the other way), well, the cargo is going to take a bit more than 15 months for the trip. Which is a long time for the cargo spacecraft to be idle, doing nothing but surrounding the cargo. It is hard to amortize the cost of the spacecraft and spacecraft maintenance when it is only doing billable work at the start and end of the trip while the engines are burning.
A few innovative thinkers had the bright idea that since the expensive engines (specifically the propulsion bus) are only needed at the start and end of the journey, why not jettison them so they can be reused? Have the cargo in cannisters or tied to a frame, and the propulsion bus is only present at the start and the destination. The cargo cannisters will probably be sized to be compatible with standard cargo cannister form factors.
I've found three techniques using temporary engines:
In this Inert Cargo Vessels scheme the propulsion bus becomes a space tug. This latches onto a cargo container, pushes (or pulls) the container into the desired trajectory, detaches to let the cargo go on its merry way, then the tug flies back to the cargo staging area for a propellant re-fill and to grab the next scheduled cargo cannister.
The cargo cannister flies in its trajectory, with no engine but no need for an engine either.
15 months later the launched cannister approaches Terra where it is intercepted by a Terran space tug. It then decelerates into the cargo storage orbit, parks the cargo, refills, and heads out to catch the next incoming cargo cannister.
Since the tugs are constantly working they can amortize their little hearts out.
A distantly related concept is the Flyaway Engine.
This Inert Cargo Vessels scheme can be used if the delta-V requirement for the trajectory is not too excessive. The pair of space tugs can be replaced by a pair of momentum exchange tethers ping-ponging momentum energy back and forth between each other as they launch and catch cargo cannisters. With this scheme it is important that each tether in the pair launches the same amount of mass that they catch. Otherwise one or the other tether will start running out of energy and will have to be spun up again with solar power or something.
Again these are are constantly working and constantly amortizing.
Back in 1976 O'Neill had a problem which was preventing the construction of his O'Neill cylinder L5 space colonies. Their isn't enough money in the entire world to boost the required million metric tons or so of construction into orbit using rockets. Therefore O'Neill designed a lunar mining base which would dig up the required materials, cheaply catapult them into cis-Lunar space with a huge Mass drivers, and when a given load approached the L2 point it would be intercepted by a huge net-like construct picturesquely named a "catcher." No cargo spacecraft required. The power requirements of the mass driver were reduced by having the source of the materials on Luna instead of Terra, since Luna has a much milder gravity well.
Some mass driver designs have the masses of ore encased in ferromagnetic cannisters to give the mass driver's magnetic field something to grab. Others need no cannisters, instead they use ferromagnetic buckets which are halted and returned to be reused at the end of the mass driver. The ore goes flying into space toward the catcher. This saves on cargo cannisters cost.
Back in the early 1900's noted genius and mad scientist Nikola Tesla figured he could tap a conductive layer in Terra's upper atmosphere and used it to wirelessly broadcast electricity. The electricity would be held in standing waves around the entire globe, and could be tapped by machines in remote locations for electrical power. It would also make the entire upper atmosphere glow, making cities and shipping lanes happy while infuriating astronomers. Oh, and it would also work as a wireless telegraph.
While many of Tesla's devices were brilliant, this one was a total crack-pot idea. Telsa was suspicious of these new-fangled ideas about air-borne electromagnetic waves. Not to mention there was no way to send an electricity bill to the people using it.
About fifty years later science fiction writer Murray Leinster wrote a series of short stories featuring a huge device called a "landing grid." I have been unable to discover the source of Leinster's inspiration, but I suspect Telsa's Wardenclyffe Tower. As far as I have been able to determine the first of these stories was Sand Doom (1955), first of the Colonial Survey series.
Anyway a landing grid is a circular arrangement of steel girders and copper cables about half a mile high and one mile in diameter. It is set firmly into the planet's bedrock.
For a planetary colony, it supplies electrical power by tapping the electrical potential difference between the ground and the planet's ionosphere. The planet acts like a huge capacitor. One plate is the ground, the other plate is the ionosphere, and the insulating dielectric is the atmosphere in between.
Since the ionosphere is basically energized by the planet's sun it will supply electricity for as long as the sun shines. As to how much energy is available, the best I can say is "lots and lots." A certain Dr. Elizabeth Rauscher estimated that the ionosphere and magnetosphere had a potential energy of about 3 terawatts. No idea of how rapidly the energy would be replenished by the sun.
The second vital function a landing grid supplies a planetary colony is landing services. It can use technobabble tractor beams to grab a spacecraft at a range of tens of thousands of miles and gently lower it to land in the center of landing grid. Or gently lift a spacecraft from the grid up into space, releasing it several thousand miles altitude. The spacecraft does not have to spend horrific amounts of delta V to get halfway to anywhere. The inexhaustible supply of ionospheric electricity will do it for you.
The framework of girders requires about one foot of diameter for every ten miles of tractor beam range. They are typically one mile in diameter, giving the tractor beam a range of about 53,000 miles (about 6.7 Terran diameters).
When a new planetary colony is founded, the first construction crew lands in rocket-propelled vehicles (since there is no existing landing grid). Their priority is to quickly build a grid to get the colony started.
In theory, interplanetary and interstellar war was not possible in Leinster's novels. Naturally a planet would not be foolish enough to use their grid to land a hostile invasion force. And the grid was perfectly capable of attacking an enemy orbiting fleet with tractor-beam launched missiles, or even rocks for that matter. Without grid support, an invasion force trying to land troops would need lots of rockets with ugly mass ratios. The invading fleet can launch missiles and bombs, but they have limited supplies (limited to what they brought with them). The planet ain't going to run out of rocks.
And if the invaders destroy the landing grid, they will lose easy access to the surface. Worse, any invading forces actually on the planet will be stranded until a new grid can be constructed. So the invaders do not want to nuke the grid, but the grid can decimate their fleet with hypervelocity rocks.
The theory was exploded in Leinster's 1957 story The Grandfathers' War. Basically they built a space-going landing grid.
Conventional grids grab objects in space with a tractor beam and pulls it to the ground. This monster grabs the ground with a tractor beam and pushes the grid into space. Conveniently the FTL drive can operate the instant a ship (or space-going landing grid) is several planetary diameters away from the planet, so the grid does not even need any rockets. Directly into FTL drive it goes. The warlike grid travels under FTL drive then emerges into real space in orbit around the target planet. There it uses its tractor beam to land itself, instantly creating an invader-controlled grid on the surface of the hapless planet. The grid then lowers the hordes of invading troop carrier starships gently to the surface and the attack begins. The only question I have is can the space-going grid tap the target planet's ionosphere while in orbit?
Lucky for the peace of the galaxy, in Leinster's universe nobody ever copied the grid-ship idea, and it was forgotten. The idea was not used in subsequent novels.
In 1962 Walter Richmond was doing research into atmospheric electricity and invented what he called the Solar Tap. It was a way to access the potential energy difference between the ionosphere and the ground, but it was rather hair-raising.
You build an insulator, a pyramid shaped pile of rock about 150 meters tall. Be sure you locate the insulator well away from the magnetic poles of the planet. From the peak is shot a powerful laser beam pulse to create a conducting ionized trail all the way to the ionosphere. A titanic bolt of lightning travels down the trail to hit the insulator. There equipment does its best to harvest as much of the lightning as it can, without destroying the equipment or too much of the surrounding landscape.
As an encore, distribute the energy world-wide by using some sort of technobabble Tesla style energy broadcasting technology.
Why is it so important to site this far away from the magnetic poles? Well, the lightning bolt will create a magnetic field cross-wise to the planet's natural magnetic field. The result is to pinch the bolt and stop it after a few microseconds. Then you shoot another laser blast to created the next lightning bolt. All nice and controlled.
If the insulator is at a magnetic pole, the lightning bolt's magnetic field will be parallel to the planet's field. The bold will not be pinched. It will be permanent until the ionosphere is depleted after a week or so (an "avalanche"). In other words about 3 terawatts of power will start evaporating the continent around the magnetic pole, split the tectonic plates and start the continents moving around, create nuclear winter, destroy all civilization and cause a global extinction event.
That would be bad.
In 1967 Walt and Leigh Richmond wrote The Lost Millennium aka Shiva. The idea behind the novel was that solar taps were not only possible, they had been invented about eight thousand years ago. The reason we were unaware of this is because the idiots back then had sited the main tap at the magnetic pole in the name of maximum power harvesting, and they resolved to be very very careful not to let an avalanche start. With predictable results. Pretty much erased their entire civilization, it did.
The reason the insulator for a solar tap is about the same size and shape as the Great Pyramid of Cheops is because the latter is an insulator for a solar tap. Apparently some survivors from Atlantis built Giza a couple of thousand years after the avalanche (the pyramid that caused the avalanche was pretty much obliterated). Well away from the magnetic pole you will note. The laser firing makes a noise that sounds like "SHEEEEE!" and the returning lightning bolt makes a sound like "OPS!". So the solar tap in operation sounds like SHEEE-Ops!, SHEEE-Ops!, SHEEE-Ops!. Which is where the Cheops pyramid got its name. Cute.
The novel includes all sorts of historical anomalies harvested from tales of Atlantis, ancient astronauts, and Chariots of the Gods? The reason archaeologists are not constantly stumbling over eight thousand year old automobiles and skyscraper girders is because the broadcast power system made large metal objects a dangerous idea.
Anyway the other item relevant to our interests is that the solar tap could also be used to boost and land spacecraft. The Richmonds are vague in the details but they maintain that a network of smaller pyramids can create a pattern of laser beams to craft a titanic Jacob's Ladder. The high-voltage traveling arc boosts or land spacecraft by electromagnetic induction. Somehow (the details are left as an exercise for the reader). In the novel, during boost mode the solar tap sounds like ANGOR-WATT! ANGOR-WATT! which is also cute.
In their later novel Gallagher's Glacier the Richmonds take up planetary liberation by solar tap. In the novel, all the poor planetary colonies are controlled by an evil corporation. The colonies are not allowed to have solar taps, because the corporation do not want the colonies to be anywhere near being self-sufficient.
Gallagher takes a tip from Leinster and mounts the solar tap on a spaceship. It is impossible for a colony to covertly build a solar tap over a couple of decades without the evil corporation goons noticing. But once Gallagher's space ship shows up, the colony instantly has a solar tap, and can use its energy to defeat the goons and kickstart building their own permanent solar tap. Corporation Revolutionary War soon follows.
3-D printing is also known as "additive manufacturing". This is because the object is created by adding blobs of new material, instead of the conventional method of starting with a block of material and carving away the unwanted bits (for example, as done by a CNC router).
This was a mind-blowing concept when Keith Laumer used it in his 1981 novel Star Colony, but with the advent of hobbyist 3-D printers it is now considered trendy but not impossibly futuristic.
Corporations will be angered by 3D printers: if you thought the RIAA went ballistic about digital music piracy and the MPAA was freaking out about movie file sharing, you ain't seen nuthin' yet. Manufacturers are going to start foaming at the mouth about digital object piracy. I predict even more draconian Digital rights management laws.
There is already in the real world people who are stirring up trouble by making blueprints that will 3D print plastic "ghost" firearms with no serial numbers. They are trying to strike a blow for Libertarianism, but they might just wind up making 3D printers illegal. Angry corporations are one thing, angry governments are even worse.
But I digress.
NASA is interested in 3-D printing because Every Gram Counts. It would be a valuable savings in mass if a spacecraft did not have to carry spare parts for every conceivable thing that might break, but could instead only carry a 3-D printer and the raw material. You do not have to waste payload on spare parts you might never need. And the computer blueprints have zero mass.
Most currently available 3-D printers only print with one material (generally some kind of plastic). Innovators are frantically working on printers that can handle multiple materials. This is vital for printing, say, an electric motor or an electronic circuit. Currently available printers deposit blobs of material, in the future they will deposit on an atom-by-atom basis.
In September 2014, SpaceX delivered the first zero-gravity 3-D printer to the International Space Station (ISS). On December 19, 2014, NASA emailed CAD drawings for a socket wrench to astronauts aboard the ISS, who then printed the tool using its 3-D printer.
As a proof-of-concept, Markus Kayser created the Solar Sinter. He noted that in the deserts of Terra, there is a lack of useful artifacts but unlimited amounts of sunlight and sand. The Solar Sinter is a computer controlled magnifying glass that 3-D prints by melting layers of sand. There are many planets and moons where such a tool would be incredibly useful.
Architecture Et Cetera (A-ETC) is working on Project SinterHab. This will use microwaves to fuse Lunar dust in order to 3-D print habitat modules for a Lunar base.
Foster + Partners is working with the ESA to make a 3-D printed lunar base. Lunar soil is mixed with magnesium oxide to produce the material. Layers are bound by being sprayed with a binding salt in a controlled pattern. The binding salt turns the material into a stone-like solid.
A 3-D printer can also be used as the "assembler" component of a Santa Claus Machine.
The Santa Claus Machine is sort of a technological version of Aladdin's Lamp or a Cornucopia. Basically it is a Star Trek Replicator that uses in-situ resources as feedstocks. More crudely it is a mass spectrometer feeding a 3D printer.
If you are trying to set up a base or colony on a desolate moon or planet, a Santa Claus Machine could be the difference between success and failure. The less equipment and prefab base you have to bring and the more stuff you can manufacture with local resources, the better.
As with any such thing, it has two parts: a disassembler and an assembler. This is because there are two basic operations possible in the universe, analysis and synthesis. That is, breaking one large object into smaller parts, or assembly smaller parts into one larger object. The ancients called this "solve et coagula" (e.g., written on the arms of the Sabbatic Goat in the famous illustration by Eliphas Levi).
The disassembler breaks down the input material into atoms, then sorts the atoms by element and isotope. This provides the raw materials needed by the assembler.
You shovel rocks, dirt, and other regolith into the hopper of the fusion torch. The input matter is flash heated to a temperature of about 15,000 K by the awesome power of thermonuclear fusion, disassembling all the compounds into individual atoms and ionized atoms at that. You now have all the atoms separated in a plume of ultra-high temperature plasma.
There are many proposed ways of sorting the atoms into bins for each individual element and isotope. The most commonly mention method is using a mass spectrometer.
Atoms have inertia, like anything else that is matter. And like all other matter the more mass an atom has, the more inertia is has. So if the atoms are moving in one direction in a atomic beam, if you give each atom a shove to the right with a given strength push the atoms with less inertia will be nudged off course more than the atoms with more inertia. Without the push all the atoms in the beam will strike the target point. The shove with smear the target point to the right. If you nudge enough, the target will smear into a row of points, one for each element. Nudge it more and the points will separate further into points for each isotope of each element.
All you have to do is put a collection bin at each target point and they will fill up with pure isotopes. But do be careful about the bins for fissionable isotopes. Allowing a critical mass to accumulate will have unfortunate consequences.
Mass spectrometers generally use a magnetic or electrostatic field to give atomic beam a shove.
Keep in mind that what you get out depends upon what you load into the input hopper. If the asteroidal regolith you are shoveling in contains no uranium, none is going to show up in the collection bins. You might have to import isotopes that are absent in your location.
Just imagine how useful the fusion torch+mass spectrometer combo would be for recycling the mountains of trash filling up our real life land-fills. The entire blasted world is impatiently waiting for somebody to tame fusion power.
Also note that this technology makes it easy to refine uranium ore into weapons grade uranium, which will make the astromilitary and the authorities extremely nervous. Current enrichment techniques such as gas centrifuges require the resources of an entire nation the size of Iran. A fusion torch could do in your garage.
The assembler takes atoms from the disassember's output, and puts them together according to the user selected blueprint.
This will basically be advanced versions of the 3D printers and rapid prototyping machines available now. Instead of just handling one material (typically plastic) they will be capable of printing in multiple materials. They will accept as feedstock the elements and isotopes from the disassembler, and either chemically create the required compounds or just print the compounds by alternating the atoms.
Early crude versions will print blobs of paste composed of compounds created from the atom feedstocks, much as a commercial 3D printer makes objects out of molten plastic. Later advanced versions will assemble the object atom by atom.
The limits will be
- the chemical elements required from the disassembler for object currently being printed (does the local regolith have all that is necessary?)
- the availability of blueprint files for the desired object (are the blueprints illegal?)
- the speed of printing the object (if it takes ten years to print, forget it)
- the supply of fusion fuel to power the fusion torch (though with fusion you are talking gigawatts per centigrams/seconds of fuel)
Faster printers will be more expensive, because that's the way it always is.
Some blueprints will be illegal (e.g., DIY nuclear warhead) and of course will be readily available anyway from data smugglers and on the dark web. There might be illegal blueprints which on the surface look innocent, but combining part 23 of the dust precipitator plan with part 17 of the air conditioner plan creates a working submachine gun.
Needless to say the invention of a Santa Claus Machine will have a drastic effect on the economy of your civilization.
And other things too. I've already mentioned how the powers-that-be will be concerned with giving rock-rats the ability to manufacture weapons of mass destruction and refine kilogram lots of weapons grade fissionables. And I'm sure the futuristic equivalent of the MPAA and RIAA will be furious with Joe Asteroid wallpapering their habitat dome with atom-level perfect copies of the Mona Lisa. Not to mention how angry the banks will be with a device that can crank out undetectable counterfeits of coins, bills, cheques, and other legal documents.
Of course things get astronomically worse if a Santa Claus Machine can produce copies of itself. Now you've got a freaking Von Neumann self-replicating machine on your hands.
I have a feeling that Santa Claus Machines will always be under military guard, much like the beam propulsion lasers controlled by the Laser Guard. The Santa Guard will place the machine at the site of a future base/colony, and watch what is manufactured like a hawk. If a colony builder submits a blueprint for something questionable, they are liable to be apprehended by the Santa Guard and questioned.
In the far future Santa Claus Machines might be equipped with a law-abiding artificial intelligence. If the user asks it to make a nuclear warhead, the machine will refuse and call the cops.
A self-replicating machine or Von Neumann device is an independent robot that can create a duplicate of itself from materials scavenged locally. The little monsters can multiply exponentially (i.e., like cancer) so it is best you have some kind of control or kill switch on them.
They are used when you have a really big job, so you want the robot work force to scale itself up to a size suitable to the task. For example: covering the entire equator of the planet Mercury with solar power cells in only a few years. Or sending robot space probes to every planet in the entire galaxy.
Plastics are organic polymers, which means they are composed of huge chains of carbon and hydrogen molecules. The raw materials can be found in carbonaceous asteroids and in the hydrocarbon lakes of the Saturnian moon Titan.
Inside the closed ecology of a spacecraft's or base's CELSS some of the carbon and hydrogen can be diverted to brewing up some plastics. The source can be from carbon dioxide in the air or from agricultural waste.
Clothing is difficult to manufacture in microgravity, from growing the plant fibers to spinning, weaving, dying, and tailoring. All of those processes are much more difficult when things are floating around. This will limit the supply of available clothing, and make them expensive.
On the ISS, the crew wears garments made of cotton. These have a tendency to shed lint which can clog up ISS machinery and air filters. They are experimenting with Merino wool shirts and polyester shorts, which are lighter and do not shed lint.
The clothing might be treated with anti-microbial agents to make them odour resistant, since a microgravity clothes washer is so problematic that the ISS does not have one. On the ISS, clothing is worn and re-worn without washing until they get too stinky. Then they are put on the next cargo supply ship to burn up in re-entry. Actually, in microgravity, clothing does not actually touch the wearer's body as much as it does under Terra's gravity. For a crew of six, the ISS requires about 400 kilograms of clothing per year.
In classic Star Trek, the laundry renders the clothing back down to its chemical components, filters out the dirt, then refabricates the clothing. Nowadays we would think in terms of a 3D printer. Later versions of Star Trek would use unobtainium "replicators", but they have unintended consequences.
There are two basic ways to enable textiles to kill microbes. The first is to coat the fabric in a liquid solution that contains metals like silver ions; metal oxides like copper oxide; or compounds of ammonium. The other way is to impregnate the threads themselves with these kind of antimicrobial agents. Some testers said that the clothing would not stink, but it did tend to get noticeably heavier the more times it was worn. Presumably from the accumulation of perspiration and cast-off skin cells.
The ISS solution of "rely upon resupply from Terra" for the clothing problem won't work for a Mars Mission. Terra will be too far away, so you'll have to carry all the required clothing. In and effort to reduce the clothing payload NASA commissioned the UMPQUA Research Company in 2011 to produce the Advanced Microgravity Compatible Integrated Laundry System. The prototype worked on a vomit comet test flight, but UMPQUA is trying reduce the unit's water and power supply requirements.
Skirts or kilts are discouraged because [a] it is difficult to impossible to keep them in a modest position in free fall, and [b] if the decks are open gratings instead of solid floors, people on the next deck down will be treated to an up-skirt view. No panchira allowed.
NASA ISS astronauts wear clothes with lots of pockets and strips of velcro, as a handy place to carry gear.
In Larry Niven's Protector, the Belters of the asteroid belt spend most of their lives inside their space suit. They have a tendency to paint their suits in extravagant colors. One of the characters had Salvador Dali's Madonna of Port Lligat on the front of their suit. In an interesting psychological quirk, Belters also tend to be nudists when in a pressurized environment. This could also be a response to the difficulty of making clothing, or a reaction to the how expensive clothing is.
There are just some industrial applications that demand power approaching Kardashev Type I levels. Hyperpower stations will supply you with massive amounts of power (along with a massive electricity bill).
Titanic solar power stations covering huge areas on the surface of Mercury or Luna are called "Asimov Arrays", name bestowed by James Powell and Charles Pellegrino after Isaac Asimov pointed out several serious errors in their design. Such as "You do know that Mercury is not tidally braked with respect to the Sun, do you not?"
Do keep in mind that it is not mandatory for the solar cells to be mounted on Mercury, they can be orbital. You could even place them closer to the sun if you really need the power. The only problem is that light pressure will tend to push them away, Mercury's gravity can anchor them.
Alternatively, you mount lasers on copper rods and launch them from Io at Jupiter. As the rods cut the magnetic lines of force they generate electricity. This is converted into laser light and beamed back to Io. Rod is destroyed when it hits Jupiter, but so what, they are cheap.
Naturally huge arrays of fusion power plants are going to require a huge supply of fusion fuel nearby.
Antimatter distribution is administered by the Antimatter Guard because is it so much easier to misuse than mere plutonium..
Atomic Rockets need Atomic Fuel. Raw uranium or thorium ore is worthless as fuel for your nuclear thermal rocket or nuclear power reactor (the same goes for nuclear weapons). The stuff the asteroid miners haul in will have to be enriched before it can be used as fuel.
Common power reactors require enrichment from 1% to 20%, fast-neutron power reactors and nuclear thermal rockets need 20% to 85%, above that is the weapons-grade fissionables needed for nuclear weapons, SNRE-class propulsion, Orion pulse units, and nuclear salt water rockets burning 90% UTB.
Enrichment requires a sizable high-tech factory, they will have to be strategically placed around the inhabited solar system. Along with security forces provided by the astromilitary of a select group of nations, to ensure that none of the weapons-grade plutonium gets stolen (the Nuke Guard).
It is a lamentable fact that fission engine fuel elements clog up with nuclear poisons and stop working while there is still lots of fuel in them. After about 15% of the fuel is burnt (85% unburnt) the rod stops fissioning. Since is it a criminal waste of scarce fuel to throw the rod away when 85% is still yet to be used, you have to take it to a nuclear reprocessing plant. The plant will filter out all those nuclear poisons and use the unburnt fuel to make new fuel rods. A distressing by-product is lots of weapons-grade plutonium, which the Nuke Guard will also have to deal with.
Obviously reprocessing will only be needed for solid core and closed-cycle gas core NTRs. Other nuclear rockets blow the nuclear fuel out their tail pipes, burnt and unburnt.
There will have to be reprocessing plants strategically placed around the inhabited solar system, perhaps inside enrichment plants. Perhaps with a network of fuel transport ships shuttling fuel rods (fresh and spent) between plants and spaceports for convenience. Said ships will undoubtedly be a part of the Nuke Guard.
Nuclear power unfortunately produces radioactive waste. The low-level cesium-137 and strontium-90 waste has a half-life of 30 years or so (decaying to 1% of it original deadly strength in about 180 years). But the plutonium is freaking transuranic waste with a half-life of around 24,000 years (decaying to 1% of it original strength in about 144,000 years, about the time separating us from Neanderthal Man).
Where are you going to dispose of this death-metal? Pretty much every intelligent being will scream in your face "NOT IN MY BLASTED BACK YARD, YOU AIN'T!!!"
A common simplistic solution is to lob the stuff into outer space (since there currently are no back yards in space). You may have seen the concept in the scifi show Space 1999. Understand that the bit where the Space 1999 disposal site blows up and kicks the moon out of orbit is utter bovine excreta.
Granted that space is so freaking huge that it is pretty much impossible to contaminate it with glowing pollution. But the transport cost makes this solution impractical. It would be several orders of magnitude cheaper to drown the stuff in vats of computer printer ink mixed with Dom Pérignon champagne and wrap them with diamond-encrusted iPhones tied with ropes of saffron.
For this to work the cost to boost payload into orbit will have to come way down, or the cost of terrestrial disposal will have to go way up. Or both.
If the boost cost becomes reasonable, an old NASA report recommends disposal in Lunar or Solar graveyard orbits from an overall mission safely standpoint. Some sort of remote-controlled space rescue capability will be needed, in case a rocket malfunction sticks the radioactive waste rocket into the wrong orbit.
No, don't even think about trying to hurl the radwaste rockets into the Sun. In delta V terms it will be far cheaper to accelerate the waste to Solar escape velocity and into intergalactic space.
The Mars mission requires hydroponics for food and air for the astronauts, nine months worth. The habitat module. And worst of all, the massive anti-radiation storm cellar. All of this takes mass. Then you have to add the mass for the lander and the other equipment you'll need on Mars. Just think about the propellant bill.
Then if you have a second expedition, you have to pay for it all again. And for each subsequent expedition.
About this time, astronautics experts had the thought "what if we could re-use some of the required equipment?" More specifically, re-use the delta-V.
Take the habitat module, the hydroponics, and the storm cellar and make it into a space station. Spend enough propellant to delta-V it up into an orbit that passes by Mars and eventually returns to Earth. It will regularly pass by Earth and Mars for the rest of eternity, with a little mid-course correction now and then. So you now have a habitat module delta-Ved for a Mars mission that can be re-used. It is a Cycler.
For your next Mars mission, you have a transfer vehicle that will carry the crew and mission specific payload. It rendezvous with the cycler, more or less paying the same delta-V cost as the start of a Mars mission. Except it only pays the propellant cost for the crew and the mission payload, it does not have to pay for the habitat module. You will be re-using the delta-V for the hab module by using the cycler. When the cycler passes by Mars, the transfer vehicle leaves the cycler and burns enough propellant (or aerobrakes in the Martian atmosphere) to delta-V into Mars orbit. The cycler goes on its merry way, still full of delta-V, still available for re-use by a future expedition.
Keep in mind that you still need the propellant for the people and mission payload. But saving the propellant needed for the habitat module is a huge help.
Understand that since a cycler is a clever way to reuse the delta V of the habitat module, the hydroponics, and the storm cellar, the implication is that a cycler is worthless for sending inert payloads to Mars. It will take the exact same amount of delta-V to send the inert payload to Mars regardless of whether you use the cycler or not, so what's the point? This is because inert payloads do not need habitat modules, hydroponics, nor storm cellars.
Hop David has computed the orbits for Earth-Asteroid cyclers, discovering the existence of virtual "railroad towns".
This is from Ion engine propelled Earth-Mars cycler with nuclear thermal propelled transfer vehicle. It is a preliminary study by California University School of Engineering and Applied Science.
The report makes a few assumptions:
- There is a space station in LEO to be a base for construction of the cycler, and a rendezvous spot for the "taxi" (spacecraft that ferries astronauts to and from the cycler)
- There is some kind of transportation system between Terra and the space station (a Space Shuttle or Soyuz spacecraft)
- Previous missions has already established habitats on the Martian surface, as well as landing/launch pads for the taxi
- Previous missions has already established an in-situ resource utilization plant to produce liquid hydrogen propellant for the NTR taxi. The cycler cannot a transport all the propellant the taxi needs, it has to refuel on Mars.
- Previous missions have already established a fuel ship capable of transporting liquid hydrogen from the Martian ISRU plant to the taxi in low Mars orbit
The heart of a cycler system is the Cycle; that is, the orbit it follows.
The study looked at various orbits, trying to optimize for:
- More frequent encounters between Mars and Terra
- Smaller detal V angles and Terra and Mars approaches
- Shorter stay time on the cycler
- Easy predictability of the position of the cycler
Having just one cycler and rotating its orbit to meet the two planets seems attractive, but there are major drawbacks. The fuel required to rotate the orbit are expensive, about 85% the mass of the cycler. This requires constant refueling. Also since the cycler is not in a predictable orbit the motion will have to be constantly monitored and mid-course corrections applied. With no corrections the orbit error will propagate to future trips.
To deal with the predictability problem a proposed solution was to rotate the orbit every 2.143 years (the delay between times the relative positions of Mars and Terra repeat) by 51.429 degrees (360° / 7, giving 7 discrete orbits). This would cover the entire range with only seven passes. The orbits would be rotated by a burn performed at the closest approach to Terra in order to get a gravity assist. The drawback is that the fuel requirements would be about the same, and there would be periods of more than ten years before the cycler returned to Mars and allowed the Mars explorers to return to Terra.
So this option was rejected.
The Up/Down Escalator orbit was ruled out because: the Taxi would need excessive amounts of delta V to catch a ride and the orbit would have to go way further past Mars in order to encounter Mars on the inward swing (which drastically increases the cycler orbital period).
This option was rejected as well.
The report concluded that the optimum cycle was using three cyclers with VISIT-like orbits. One at zero degrees, one at +130° and one at -130° (230°). This allows squeezing the most mission into each 20 year period while optimizing the other factors.
In Table 2, row Cycler DELTA V (row 19) shows the taxi delta V needed to leave the cycler and enter close Mars orbit. This varies from 5.27 to 6.32 kilometers per second. Row Hyperbolic delta v (27) shows the taxi delta V needed to leave the cyclear and enter close Terra orbit. This varies from 9.49 to 10.46 km/s. They figure that to perform these maneuver the taxi will need a thrust of 5.639e+5 Newtons, which is good because the planned taxi engine will have a thrust of 6.98e+5N.
|ORBIT 1 Mars Approach|
|Burn Time||378 s (6.3 min)|
|ORBIT 1 Terra Approach|
|Burn Time||432.7 s (8.05 min)|
|ORBIT 2 & 3 Mars Approach|
|Burn Time||333.78 s (5.56 min)|
|ORBIT 2 & 3 Terra Approach|
|Burn Time||507.3 s (8.46 min)|
Table 3 shows three Terra-Mars mission opportunities over an 18 year period. Trip 1 starts at year Zero, uses the zero degree cycler, and has a mission duration of 5.41 years. Trip 2 starts at year 4.75 and has a mission duration of 6.73 years. Trip 3 starts at year 14.89 and has a mission duration of 2.856 years.
The Reactor produces 10 MWe (electrical) power. The Power Conversion is a Stirling cycle with an efficiency of 0.254 so the reactor has to produce 40 MWth (heat). To reject waste heat 1,000 m2 of Heat Radiators operating at 1,000K are used.
The Storage / Experiment / Greenhouse module is above the hub. It contains space for microgravity experiments, food storage, and the life support reclamation systems. The hygiene/gray water reclamation system uses various filters, as well as Waldman's dark green lettuce in the greenhouse.
The cycler provides artificial gravity by spinning as a dependent centrifuge, where the spin axis is parallel to the thrust axis. The habitat modules are set at a radius of 50 meters from the spin axis, the spin rate is 2.32 rpm (at the nausea limit for the untrained), the resulting gravity is 0.3 g. The report assumes this will be enough gravity to prevent muscle atrophy, since it will be very hard to explore Mars if the astronauts are too weak to walk. For what it is worth the surface gravity of Mars is 0.376 g.
The bulk of the cycler's mass is on the spin axis (with the exception of the habitat modules) to make the smallest moment of inertia. This reduces the amount of reaction control jet fuel needed to spin up or spin down the cycler. The jets are located at the ends of each habitat module. The cycler must be despun for docking and releasing the taxi.
Each of the two Habitat Modules has two levels: command level above and residential level below.
The command level has the control/communication room, the kitchen, the communal room (including exercise equipment), and the infirmary. On such a multi-year mission a dedicated sickbay is needed. And two of the sixteen crew are doctors.
The residential level has the crew quarters, toilets and showers. Each crew member has their very own 2 x 3 meter private room, with bed and desk.
The Storm Cellar is located below the hub. It can hold all sixteen astronauts and is designed to ensure that a solar proton storm does not inflict a dose higher than 0.5 Sieverts. The astronauts are seated in semi-reclined chairs so they can sleep or do work, since the ceiling is too low to stand up (2.5 meters floor to ceiling).
Aluminum shielding instead of water was chosen due to ease of construction and maintenance. The shielding is 20 grams per square centimeter of aluminum (thickness of 7.4 centimeters). The largest proton storm ever recorded was the August 1972 solar event and the most harmful spectrum was the February 1956 solar event. If the cycler suffers a solar event with the duration and intensity of the 1972 event coupled with the deadly spectrum of the 1956 event the storm cellar will ensure the astronauts only suffer a dose of 0.43 SV.
The storm cellar has an expensive mass of 15,000 kilograms, but radiation shielding always has a painful amount of penalty mass.
The Communication Array is de-spun. It is mounted on a coupling with rings and brushes (perhaps this could be replaced by a Canfield Joint).
The Ion Thrusters use argon propellant. Each engine has a mass of 165 kg, a diameter of 0.85 m, a specific impulse of 10,000 sec, a propellant mass flow of 1.579E-3 kg, and produce 4.4 Newtons of thrust. The reactor produces 10 MW of electricity but for safety and to leave power for the rest of the ship only 9.5 MW are used by the engines. For the 130,000 kg cycler, 35 ion engines were deemed adequate. Since ion engines tend to fail, 60 engines are carried.
|Burn Time||8,356,554.78 s (96 days 17 hours 15.9 minutes)|
|ORBIT 2 and 3|
|Burn Time||9,077,264.09 s (105 days 1 hour 27.7 minutes)|
|Mass Flow||42.6 kg/s|
|Dry Mass||11.65 m/s2|
|ORBIT 1 Mars Approach|
|Burn Time||378 s (6.3 min)|
|ORBIT 1 Terra Approach|
|Burn Time||432.7 s (8.05 min)|
|ORBIT 2 & 3 Mars Approach|
|Burn Time||333.78 s (5.56 min)|
|ORBIT 2 & 3 Terra Approach|
|Burn Time||507.3 s (8.46 min)|
The taxi has two stages: a nuclear powered first stage with a solid core NTR and a chemical powered second stage (based on a McDonnell Douglas DC-X).
Contrary to what you'd expect, the nuclear stage never lands on Mars. Its purpose is to ferry the chemical stage from the cycler (as is whizzes past Mars) to low Mars orbit. The chemical stage separates and lands on Mars while the nuclear stage stays in Mars parking orbit. It seems that the designers were hesitant to bath a part of the Martian surface with deadly radiation from the nuclear engine. It was also a challenge to protect the astronauts from getting a bad dose of radiation as they crawled down the ladder along the taxi's side to step on the Martian surface.
While the astronauts explore the Martian surface, a robot propellant transport containing a full load of ISRU liquid hydrogen blast off and refuels the nuclearr stage in parking orbit. The chemical stage on the surface also has its liquid hydrogen tanks topped off.
When the return cycler approaches, the astronauts blast off from Mars in the chemical stage, rendezvous with the nuclear stage, and use the nuclear engine to rendezvous with the cycler.
The report is very vague on the chemical stage, other than stating it has a maximum mass of 8,386 kg. Doing some back of the envelope calculations I figure it will need about 3,550 m/s of delta V to land or blast off from Mars. Using LH2/LOX chemical engines with 4,905 m/s of exhaust velocity, the chemical stage will need a mass ratio of 2.06 in order to produce enough delta V. If the wet mass is 8,386 kg, then the propellant mass is 4,315 kg and the dry mass is 4,071 kg.
I find the figures for the Taxi Requirements puzzling. It says the Orbit 2 Terra Approach burn requires 21,613 kg of propellant. However, if the total wet mass is 30,000 kg, minus the 8,386 kg for the second stage gives us a wet mass for the first stage of … 21,613 kg. Subtact the required propellant from that and you discover the dry mass of the first stage is zero. Which is impossible. I am re-reading the report to try and figure this out. It could be that they are assuming that at the Terra Approach Burn the chemical stage will have empty fuel tanks.