The improvised space warcraft are the type that seems to hold the most story potential.  These would, as mentioned, likely be built by colonies that are in conflict.  As they do not have to operate in an atmosphere, and are built by relatively poor colonies, they are likely to be rather crude.  The basic components required are structure, propulsion, weapons, life support, power, sensors, control, and communications, and each will be briefly discussed in turn.

There are two methods of assembling an improvised warcraft, either adapting an existing vessel, or constructing a new one from parts.  The use of an existing vessel removes the need for some, though not all, of the various components.  The structure will obviously be preserved, and propulsion and life support are almost certain to remain unchanged as well.  Weapons would obviously be added, along with their associated control equipment.  Existing power supplies might or might not be sufficient, depending on the weapons fit chosen.  Sensors and communications are gray areas, depending on the tasks required of the vessel, and the existing fits in these areas.

Structure is one of the easiest components to create.  So long as the builder does not mind the craft being somewhat heavy, slapping a few beams together should be sufficient.  Existing structures, such as cargo containers, could easily be modified, or simple new ones fabricated.  In any case, this is not likely to be a driving factor in construction or conversion of vessels.  Any group incapable of creating basic structure is also almost certainly incapable of surviving if it were to win a rebellion.

Propulsion for improvised craft is likely to be chemfuel due to the fact that it is by far the simplest to implement, and has sufficient delta-V for any operations that do not involve transiting deep space.  It is entirely possible that a colony will have standard chemfuel engines used in various places, and one of them, with appropriate fuel tanks, would be fitted to the vessel.  Nuclear propulsion is much more expensive, and might well involve detailed engineering to avoid killing the crew.  The performance advantage over chemfuel for nuclear-thermal is probably not significant enough to justify the problems involved, and nuclear-electric is only practical for vessels intended for deep space use.  

Weapons are a tricky issue. These are likely to be improvised as well, and would fall into the same categories already discussed.  Improvised lasers are highly unlikely.  Industrial lasers lack the optics trains required for weapon use, while any optical trains available (probably from astronomical or other scientific sources) are unlikely to be able to handle the high powers output by the lasers.  With some time, an appropriate optical train could be designed and mated to an industrial laser, and it is even possible that colonies could design and test such things in case of war.  Kinetic projectors are in largely the same boat.  While small mass drivers could be adapted to such a task, it is difficult to see a role for such devices on a colony.  There is also the issue of targeting, which, while not insoluble, requires good pointing accuracy and possibly the creation of guided projectiles, which have even less peacetime use then the launchers themselves.  

This leaves three options, missiles, lancers, and unguided kinetics.  Unguided kinetics can be as simple as junk thrown out of the airlock, but they are of very limited effectiveness, as shown in Section 8.  Missiles and lancer projectiles face many of the same issues, and the only practical difference is the motor, which should be relatively easy to improvise.  A missile or lancer will require sensors, thrusters, and guidance logic of its own.  As this force is presumably facing another more or less improvised one, complicated guidance logic is probably unnecessary, and proportional navigation is quite easy to implement.  The sensor might well be adapted from another role, which means that the likely problem is in the thrusters.  These must be well-balanced and integrated with the guidance logic.  Depending upon the materials available, this could range from very easy (if there are large numbers of small, self-propelled objects that can be used as warheads lying around) to extremely difficult (if the entire system must be designed from scratch).  Small thrusters themselves are an unknown.  There are some systems that might use small thrusters, such as thruster packs for spacewalkers, in which case the actual integration is all that is required.  However, the number available might well be strictly limited, forcing the builder to start from scratch.  Note that this is not as easy as it seems.  While a primitive kinetic could probably be built with the simplest of systems, it would be inefficient, of dubious reliability, and probably quite large.  In the end, this is an area in which the specific situation plays a very large role, leaving us unable to anticipate exactly what might occur.  

Life support should be straightforward to build into a vessel.  Any space colony will undoubtedly have small, portable habs that can be used for surface expeditions or what have you.  Mounting one of those would be relatively simple, and the actual mechanisms for short-term life support are fairly rudimentary, easing implementation if for some reason a hab had to be constructed from scratch.

Power is a rather tricky proposition.  Unless a nuclear propulsion system is used, power is likely to be at a premium.  Most non-nuclear power studies assume that solar panels will be used, but these have significant drawbacks for space warfare.  The biggest problem is that solar panels are vulnerable to damage from opponent’s lasers or powder weapons, and cannot be angled for protection, unlike radiators.  Radiators, discussed in Section 7, are both somewhat less vulnerable to damage, and can be kept edge-on to the enemy.  A clever opponent could manage to create a dilemma between getting power and preserving the panels from damage.  Alternatives include fuel cells and batteries.  Fuel cells are the current solution for short-duration spacecraft, due to their possessing higher specific energy than batteries.  The problem is that fuel cells are somewhat involved to manufacture and are not likely to be common on space colonies, unless the colony is far enough from the Sun that solar panels are not effective.  Batteries are somewhat more likely to exist, but are heavy for their power output.  The only bright side is that a truly improvised spacecraft is unlikely to need much in the way of power, particularly when compared to a nuclear-electric laserstar.

Sensors are probably the biggest unknown.  A proper space warcraft needs some form of active sensor to localize the enemy, although it is possible that a simple passive sensor would be adequate for simple missiles.  The sensors might or might not be readily available.  Sensor suites for existing spacecraft are the most likely source, although cobbling sensor suites together from other uses might be possible.

Control is mostly a matter of systems integration.  Depending on the nature of the systems involved, control setup could range from simple running of cables and plugging together a few modules, to having to write all of the code to make everything talk, or simply doing without an integrated control system.  While it is certain that some systems will have to talk (sensors and weapons spring to mind) a large portion of the integration could be skipped, with a resulting loss of efficiency due to the crew having to move around.

Communications is fairly simple, as one thing people in space will have to do is talk.  This should ensure the availability of communications modules, which can be attached to the vessel.  The most likely cause of problems is lack of strong encryption and particularly electronic warfare capability in such modules.  Depending upon the capability of the opposition, this may or may not be a serious hindrance.  The encryption capabilities should be a reasonably simple fix, involving mostly software updates.  The EW work will be harder, as there are likely to be physical changes required to ensure freedom from interference by enemy radio transmissions.  This is not likely to be a problem if the communications module is radio-based.

This is the living breathing core of all rocket design. Delta Vee equals Vee Ee times Natural Log of Arr. This is the secret that makes rocket design possible. Now it is time to see the practical application of the key to rocketry.

Listen up, rocket designers. Write these words in letters of fire on your cerebellum. Every Gram Counts! Add an extra gram and you will pay for it with extra propellant as if the Mafia loan shark wants you to pay up with liquid hydrogen. With interest compounded hourly.

The only exception is if you are dealing with an Orion drive spaceship or other torchship.

(ed note: Captain Randall, his crew, and his submarine were inadvertently put into suspended animation in 1945. They have been awakened one thousand years later, and must go into battle with alien invaders. Captain Randall is talking over the refurbishment of his submarine with a robot named Austin. The discussion is about limitations in submarine design, but the "everything is connected to everything" principle still holds true.)

     "I'll take your word for it, Austin. At the moment there are more important things. What is the state of the ship?"
     "As on the day of trials, sir, without fault, and completely refitted. The experts have many ideas for improvements, sir, which will be submitted to you in due course."
     "And the munitions?"
     "None, sir. They had deteriorated beyond safe use. However, they can be replaced or, if you prefer it, something more violent substituted."
     "One thing at a time. At the moment, have them replaced. Does this age understand explosives?"
     "Chemical history does, sir. Your ammunition holds can be filled in approximately half an hour and that includes the torpedo tubes. Will you continue with ancient compressed air, underwater missiles, sir? We have some smaller and far more efficient."
     "Normal torpedoes with a normal warhead, please, Austin."
     "As you wish, sir."
     Randall smiled faintly. "You sound disappointed, Austin. So, since you have a logical mind, I will be logical. You probably have in mind more advanced and far more destructive missiles which came into being just before man decided on peace for all time. No doubt these missiles are faster, have a greater range, are self-seeking and are several times more destructive. How big are they, Austin?"
     "About the quarter of the size of yours sir. I estimate you can carry—"
     "They won't fit our torpedo tubes, will they? This will mean interior adaptation, rebuilding, all of which will have to be related to buoyancy, speed—above and below surface—diving and surfacing. Which method is going to be the quickest?"
     "I had overlooked that angle, sir." There was considerable respect in the robot's voice.
     "That is because you are not yet a sailor—and let me assure you. You cannot become a seaman from text books or even convenient memory banks: It's a question of experience. No doubt, Ordinary Seaman Austin, your mind is brimming with ideas for replacing the ship's motors with something modern, smaller, a hundred times more efficient and one tenth the weight. This would play hell with our diving, upset our displacement and require the alteration of our tank capacity to meet our increased buoyancy. Am I making myself plain?"
     "Too plain, sir. With respect, I am learning a great deal."
     "Good. Don't misunderstand me, Austin. I want your ideas and I shall ask for them, but only when you, yourself, can relate them to experience."

• 

I will argue that deep space craft have essentially two sections that can largely be treated separately from one another. One section is the propulsion bus — drive engine, reactor if any, solar wings or radiator fins, propellant tankage, and a keel structure to hold it all together. The other is the payload section that it pushes along from world to world.

There are both conceptual and economic reasons to treat them separately. Conceptually, because a propulsion bus might push many different payloads for different missions, such as light payloads on fast orbits versus heavy payloads on slow orbits. A little noticed but important feature of deep space craft is that you cannot overload them. They do not sink, or crash at the end of the runway, or even bottom out their suspension. They merely perform more sluggishly, with reduced acceleration and (for a given propellant supply) less delta v.

---

Conceptual logic is also economic logic. The outfits that build drive buses would like to sell them to lots of different customers for a broad range of assignments.

This is not necessarily an argument for true modular construction, with drive buses hitching up to payloads on an ad hoc basis like big-rig trucks and trailers. Building things to couple and uncouple adds complexity, mass, and cost — plug connectors, docking collars, and so forth. Moreover, drive buses intended for manned ships need to be human-rated, not just with higher safety factors but provision for supplying housekeeping power to the hab, etc. But these things, along with differing sizes or number of propellant tanks, and so forth, can all be minor variations in a drive bus design family.

The payload we are most interested in is, naturally, us. The main habitat section of a deep space ship closely resembles a space station. It is likely that habs intended for prolonged missions will be spun, for health, efficiency, and all round convenience. (Flush!) The design of a spin hab is dominated by the spin structure and — unless you spin the entire ship — the coupling between the spin and nonspin sections.

Because ships' spin habs have the features of stations they may be used as stations, and again we can imagine design families, with some variants intended for ships and others as orbital platforms having only stationkeeping propulsion. Habs are the one major part of a deep space ship that correspond fairly well to our concept of a hull. Spin habs are entirely different in shape, but the shape is constrained; once you build it you can't easily modify it, beyond adding another complete spin section.

---

For those with bank cards at the ready, buying a deep space ship might be not unlike buying a computer. If your mission needs are fairly standard, you check off options on a menu. Those with more specialized requirements can select major components — perhaps a drive bus from one manufacturer, a main crew hab from another, along with custom payload sections, service bays, and so forth, assembled to your specifications.

In fact, both technology and probable historical development suggest that fabrication and overall assembly will be two distinct phases, carried on in different places, quite unlike either shipyard or aircraft assembly practice. In the early days, large deep space craft will be built the way the ISS was, assembled on orbit out of modules built on Earth and launched as payloads. In time fabrication may move to the Moon, or wherever else, but final assembly (at least of larger craft) will continue to be done at orbital facilities. I call them cageworks, on the assumption that a cage or cradle structure provides handy anchoring points for equipment.

For game or sim purposes, my advice would be to treat drive buses and hab sections as the primary building blocks for ships, whether these components are permanently attached to each other or simply coupled together. Both approaches might be in use.

A couple of provisos. All of the above applies mainly to deep space craft, especially with high specific impulse drives. Ships for landing on airless planets have some similar features. Ships that use rapid aerobraking, however, are aerospace craft and broadly resemble airplanes, even if they never land or even go below orbital speed.

And I have said nothing of warcraft. Kinetics are essentially just another payload. Lasers, and other energy weapons such as coilguns, probably draw power from the drive reactor, calling for some modifications in the drive bus. These things don't much affect the overall configuration. Armor protection would, but discussions here have left me doubtful of its value against either lasers or kinetics. Laser stars and other major warcraft may not be dramatically different in appearance from civil craft of similar size.

---

• 

Nyrath: I'm reminded of Sir Arthur C. Clarke's early space science books. He noted that a nuclear powered spacecraft would probably resemble a dumb bell, that is, two spheres connected by a stick. The hab module is the forward sphere, the nuclear drive is the rear sphere, the stick is long to provide some inverse square protection from radiation, and the propellant tanks would be on the stick, probably clustered near the nuclear drive.

One can also imagine modules designed by diverse corporations being incompatible with others on purpose. "Not invented here" syndrome.

One can also imagine a tramp freighter composed of incompatible modules, being held together with bailing wire and spit.

---

• 

Qwert: The compatibility between modules will mostly depend on how the market develops. One extreme maybe Microsoft, a monopoly that basically sets the standard, as everything has to adhere to it. The other one extreme is the current hardware industry. Your memories and components have to fit on every motherboard if you want to sell them.

On the field of big aircraft manufacturing, standardisation dominates almost everything... excerpt the end product. If you manufacture engines, they have to fit on Airbus as well as on Boeing's aircraft. The components industry is dispersed and competition is intense.

On the other side, a pilot trained to flight with a Boeing can´t immediately switch to an Airbus without some training. Competition centers around two big players and nobody is interested in making life easier to the other.

In short: how spaceship components will be build, will mostly depend on how the industry evolves. A monopoly, strict government regulation, competition between many small producers or a highly dispersed specialised components industry may benefit a system of standards. On the other side competition between a small number of giants, may produce different incompatible systems.

---

• 

Rick Robinson: True modularity is by no means a given. But some features of modularity, call it demi-modularity, are inherent to deep space technology.

You probably want to keep your propellant tanks separate from the corrosive, explosive stuff we breathe. Drive engines are essentially bolted onto the tail. Generally the major parts of a deep space ship don't have to fit together snugly. If you want to hang something out on a bracket you probably can.

---

Ferrell: I think that the engine package, mated to a suitable tank, mated to a hab module, mated to a mission module (a seperate entity from the hab module) would (as the last step in design) incorporate the heat management system suitable to the final design. Then construction/assymbly would occure.

If there aren't landers/shuttles at the final destination and you need them, then you can carry landers/shuttles in place of cargo. A mission module may be an extended docking module that a number of small modules 'plug' into, (or your transfer craft).

Manuvering thrusters should be at mutliple points/modules (distributed from the 'nose' to the 'tail').

Of course, you could stand it all on it's ear and have the mission module be on the inside of the ship, the hab ring be around the middle (with its radiators in arcs between its connecting pylons),with the engines, tanks, powerplants, radiators, and nav sensors clustered around both ends of the mission module; and any docking would be at the tips of the mission module, or on the inside walls of the mission module.

---

• 

Nyrath: There might be a brisk trade in "interface modules", that would connect modules made by different manufacturers. I'm reminded of the Apollo ASTP Docking Module used in the Apollo-Soyuz mission. This was a tiny airlock module with a NASA style docking collar on one end, and a Soviet style docking collar on the other.

Not to mention the International Space Station Pressurized mating adapters.

---

Rick Robinson: Sabersonic — Yes, I'm gliding over a host of devils in the details. The payload section will surely have attitude thrusters, for example, and these must coordinate with attitude thrusters on the drive bus end.

Nyrath — 'Standards,' indeed! Again this is a promo for full service cageworks that will provide things like docking adapters.

• 

---

Jean Remy: On standards:

Currently the main type of freight vessel is the container-carrier. They are favored because the containers can be loaded/unloaded easily, then simply popped on a freight train or a truck. I can see the same things for a cargo spaceship. There won't be a cargo "module" but rather anchoring hard points for standard containers. Those containers will be loaded on a booster to orbit, transferred onto a ship, then at the other end the container is loaded into a simple remote-controlled lifting body (for planets with an atmosphere, say, Mars) or just a simple frame with thrusters (for the Moon) where they will be loaded into maglev trains if needed, etc.

The command post/bridge:

I'd still put in in the non-rotational part of the ship, certainly on warships (you want it as deep inside the ship as you can for protection) Not only that, but I would stop rotation in combat: precession might alter your maneuverability, and damage might weaken the structure or cause a wobble which would rip your ship apart. I would also put it in the middle for a civilian vessel, as then the command module can double as storm cellar, or if there's a separate storm cellar, you'll still have access to the command post during a solar flare. It's also a good defense against micrometeorites. The command center is just too important to risk placing it on the outer rim of a ship, no matter how convenient or comfortable it would be to the crew.

---

• 

Sabersonic: Even so, it would probably be prudent to not have all command, or at least navigation control, be monopolized by the CIC. At minimum and barring mass budgeting, there should be two for overall spacecraft control: One that is the CIC/Navigation primary with the second being the engine room for mostly emergency purposes. If only because for something as complex and (for the first few decades if not centuries of interplanetary travel, let alone interstellar) inevitably as fragile as a spacecraft, one should avoid "putting ones eggs into one basket" and to always "have a plan B". Space and interplanetary/interstellar travel is not a place for the ill-prudent, and that's just the natural dangers.

As for standards, well, before standards could ever become "standard" for lack of a better word, there lies the inevitable "Format War" in one form or another that would occur when one believes that their system is more efficient and reliable than the other or worse: a new industry/business standard that has the potential to supplant the jobs of numerous Dock workers or in this case "Cage Workers" that could drive hostility and perhaps a little bit of political pressure before the whole matter is settled in one way or another. It would not be a full parallel to the Teamsters Strike that is basically Teamsters vs Trucks, but it would not be a completely quiet deal when paychecks are involved.

---

• 

Rick Robinson: My division of ships into drive buses and payload sections is more because of operational factors than manufacturing considerations.

For example, ships inherently have (at least) two big 'hull' structures, the crew hab and the main propellant tank. These are probably at very different temperatures, which right away is a big reason to keep them physically separate.

Using the propellant tankage as shielding is appealing, but even with a vacuum separation it means that your 290 K hab shell is dumping 350 W/m2 of waste heat right into your 20 K liquid hydrogen tanks.

Jean — The containership principle seems highly likely for most space cargo, with the standard pod being defined originally to fit Earth orbital shuttle bays.

In this case, your typical space freighter is a drive bus pushing a rack structure with clamps for pods. The rack might be configurable so that you can also carry 'oversize' loads.

Jean and Sabersonic — In a parallel discussion at SFConsim-l, the question was raised whether civil ships need a 'control room' at all, or whether people could just stand watch from their regular work stations.

I think that any spacecraft with a fair number of passengers or crew will have a watch at the main life support panel, because life support is always running, has constantly changing loads, and things can go very bad very quickly. The life support panel will almost certainly be in the spin section, because that is where the life support is.

You may as well put the engineering panel here as well. There's no reason for an 'engine room' in the maritime sense, since the drive is mounted externally, and if it is nuclear you don't want to go anywhere near it.

No doubt you could maneuver and navigate the ship from here as well. But en route there is very little of this to do.

The only time you really conn a (civil) spacecraft is during rendezvous and docking, or similar evolutions. At these times you surely de-spin, but you might want a separate control station next to the main airlock, with viewports for maximum situational awareness.

Because of its location, this station would also naturally serve as the ceremonial 'quarterdeck' where VIPs are greeted, and ordinary mortals report aboard.

Warcraft are a whole 'nother matter, with protective considerations arguing for a control room at the center of the hab section.

---

• 

Jean Remy: Multiple Command posts: Yes. That's why I used the term command post rather than CIC, Bridge, DCC, Main Engineering. It was meant to be generic and refer to all these. There ought to be at least CIC and DCC, if not a redundant third post, all of them capable of doing every job, but each ideally suited to one. The ship is more efficient with all of them, but can still work with the loss of one or two.

Decentralized Command posts: Possibly, but I don't think it likely. First of all we could do that now, but we don't. Psychologically, I think crews want to have the "Captain on Bridge" so to speak. It is generally recognized that good officers are the ones who stick by their men when the going gets tough. Caesar rode into battle at Alesia, rallying his troops to victory when it looked uncertain. Washington rode with his men. Patton led from a tank. Now granted the captain of a ship can't go anywhere, but seeing him on the Bridge will be a boost to morale.

---

Rick Robinson: The question of having a control room at all was in the context of civil spacecraft. If they have an sudden emergency it is most likely to be a life support crisis such as fire, for which the classic 'bridge' functions are fairly irrelevant.

Military craft are a special case, and I'd certainly expect them to have a control center.

And maybe all ships, because to extend a point Jean makes, existing space programs are quasi-military in origin. The military outlook toward emergency response is coded into their DNA, so to speak.

All human carrying deep space ships will need a storm shelter in any case, and it would be fairly natural to configure this as an emergency control center.

---

• 

Luke Campbell: Re: Shielding. It now seems likely that a plasma magnet generated by a low mass antenna could deflect any charged solar radiation, so the crew would be safe from flares and CMEs. It does not seem like a plasma magnet could stop galactic cosmic rays, GCRs are a steady source of background radiation, not the sort of thing that a "storm cellar" would help with. A plasma magnet also would not protect against neutral particle radiation, but the only neutral particle radiation likely to be a threat is man-made: neutron radiation from nukes and possibly high energy photons from x-ray or gamma ray lasers.

For those not in the know, a plasma magnet uses low frequency radio waves to produce a rotating field that induces a current in the surrounding solar wind plasma. This current forms a dipole magnetic field that deflects and reflects charged particles. The field is not strong enough to deflect solar wind protons, but it does deflect the electrons, leading to charge separation that pulls the protons back to the electron cloud before they reach the section being protected.

Yes I was trying to stop GCRs (but calculations showed it would need an unreasonably huge magnet). The plasma magnets wouldn't stop the solar wind protons either, when considered as individual particles — you need the plasma effects of the electrons to stop the solar wind. This lets you get by with a much smaller magnet.

However, there is one possible additional method of mitigating the GCR dose - medication. As we learn more about cellular repair and cell "suicide", new treatments may become possible for both chronic and acute radiation poisoning (and oddly, you are likely to want the opposite reaction in these two cases - for chronic exposure, you want the damaged cells to destroy themselves to prevent cancer; for acute exposure you want the damaged cells to repair themselves to prevent anemia, hemophilia, a compromised immune system, and digestive difficulties). Incidentally, it has been shown that vitamin D helps with chronic radiation exposure, although the mechanism is not clear.

---

• 

Rick Robinson: To keep cryogenic propellant from boiling off on long missions you will need active refrigeration, pumping heat out of the tank. Otherwise heat buildup will be inexorable.

The ship will be designed to keep the propellant tanks away from the main radiator fins and such, and generally minimize heat absorption from the rest of the ship, so the heat you mainly have to deal with is from sunlight.

The tanks will be painted white or silvery to reflect away most sunlight. I assume that you can reflect about 90 percent. For a spherical tank at 1 AU, that means about 35 W/m2 of absorbed solar radiation that you'll have to pump out of the tank.

A 20 meter diameter tank holds about 250 tons of hydrogen, or 1500 tons of methane. Surface area is 1250 m2, so at 1 AU you'd need 44 kW of refrigerating capacity, i.e. heat extraction, to keep propellant cold.

At Jupiter distance, 5 AU, solar flux is reduced by 96 percent, and you only need 2 kW of refrigeration for hydrogen - none for methane, which will stay liquid or even tend to freeze.

Suppose you put a 10 meter diameter hab inside the tank {to give the hab some radiation protection}. (This only reduces tank capacity by 1/8.) But even with a vacuum layer you will have IR heating from the hab surface, at room temperature: 400 W/m² * 314 m² = ~125 kW.

So in this case putting the hab inside the fuel tank multiplies your propellant refrigeration bill by 4x. Which is a lot, but not horrible; the shielding might be worth it. But wrapping propellant around a spin hab is tougher.

---

Rick Robinson: Modular design is always a tradeoff. You get more operational flexibility, at cost of more complicated/heavier/weaker connections. Integral designs will be favored when the components will consistently be used together.

Much will depend on tech. Torch type drives and even 'conventional' nuke electric drives pretty much have to be mounted on a pylon, which sort of invites the option of unbolting it from the rest. OTOH, as you note, the drive section may well have its own control center. And since the rest of the ship sits on top of the pylon, it's a fine line between 'pylon' and 'chassis.'

On naming, I could also make a case that the crew hab compartment is the main component, and so would be named. Especially if it is a spin gravity structure. And 'spaceships' may end up having more than one name, just as a named train might included Pullman cars with names of their own.

And if ships are highly modular, some terms might be borrowed from railroading. For example, 'consist' as a noun (pronounced CON-sist) for the whole assemblage. Thus, 'The Ty Cobb departed Mars with a consist of [such and such modules].'

Amusing side note: modular spacecraft reverse the order of trains: the 'locomotive' or drive engine is at the back (more precisely the base), while the 'caboose' or control cabin might well be at the front/top.

This section is intended to address some gaps in available information about spacecraft design in the Plausible Mid-Future (PMF), with an eye towards space warfare.  It is not a summary of such information, most of which can be found at Atomic Rockets.  The largest gap in current practice comes in the preliminary design phase.  A normal method used is to specify the fully-loaded mass of a vessel, and then work out the amounts required for remass, tanks, engine, and so on, and then figure out the payload (habitat, weapons, sensors, cargo, and so on) from there.  While there are times this is appropriate engineering practice (notably if you’re launching the spacecraft from Earth and have a fixed launch mass), in the majority of cases the payload mass should be the starting point.  The following equation can be used for such calculations:

Where P is the payload mass (any fixed masses, such as habitats, weapons, sensors, etc.), M is the loaded (wet) mass, R is the mass ratio of the rocket, T is the tank fraction (or any mass that scales with reaction mass) as a decimal ratio of such mass (e.g., 0.1 for 10% of remass), and E is any mass that scales with the overall mass of the ship, such as engines or structure, also as a decimal.  

This equation adequately describes a basic spacecraft with a single propulsion system.  It is possible to use the same equation to calculate the mass of a spacecraft with two separate propulsion systems.

The terms in this equation are identical to those in the equation above, with R₁ and T₁ representing the mass ratio and tank fraction for the (arbitrary) first engine, and R₂ and T₂ likewise for the second.  Calculate both mass ratios based on the fully-loaded spacecraft.  If both mass ratios approach 2, then the bottom of the equation will come out negative, and the spacecraft obviously cannot be built as specified.  Note that when doing delta-V calculations to get the mass ratio, each engine is assumed to expend all of its delta-V while the tanks for the other engine are still full.  In reality, the spacecraft will have more delta-V than those calculations would indicate, but solving properly for a more realistic and complicated mission profile requires numerical methods outside the scope of this paper.

One design problem that is commonly raised is the matter of artificial gravity.  In the setting under discussion, this can only be achieved by spin.  The details of this are available elsewhere, but these schemes essentially boil down to either spinning the entire spacecraft or just spinning the hab itself.  Both create significant design problems.  Spinning the spacecraft involves rating all systems for operations both in free fall and under spin, including tanks, thrusters, and plumbing.  The loads imposed by spin are likely to be significantly larger than any thrust loads, which drives up structural mass significantly.  This can be minimized by keeping things close to the spin axis, but that is likely to stretch the ship, which imposes its own structural penalties.  A spinning hab has to be connected to the rest of the spacecraft, which is not a trivial engineering problem.  The connection will have to be low-friction, transmit thrust loads, and pass power, fluids, and quite possibly people as well.  And it must work 24/5 for months.  All of this trouble with artificial gravity is required to avoid catastrophic health problems on arrival.  However, there is a potential alternative.  Medical science might someday be able to prevent the negative effects of Zero-G on the body, making the life of the spacecraft designer much easier.  

When this conclusion was put before Rob Herrick, an epidemiologist, he did not think it was feasible.

---

     “The problem is that they [the degenerative effects of zero-G] are the result of mechanical unloading and natural physiological processes. The muscles don't work as hard, and so they atrophy. The bones don't carry the same dynamic loads, so they demineralize. Both are the result of normal physiological processes whereby the body adapts to the environment, only expending what energy is necessary. The only way to treat that pharmacologically is to block those natural processes, and that opens up a really bad can of worms. All kinds of transporters would have to be knocked out, you'd have to monkey with the natural muscle processes, and God knows what else. Essentially, you're talking about chemically overriding lots of homeostasis mechanisms, and we have no idea if said overrides are reversible, or what the consequences of that would be in other tissues. My bet is bad to worse. As the whole field of endocrine disruptors is discovering, messing with natural hormonal processes is very very dangerous.

     Even if it worked with no off-target effects, you'd have major issues. Body development would be all kinds of screwed up, so it's not something you'd want to do for children or young adults. Since peak bone mass is not accrued until early twenties, a lot of your recruits would be in a window where they're supposed to still be growing, and you're chemically blocking that. Similarly, would you have issues with obesity? If your musculature is not functioning normally (to prevent atrophy), how will that effect the body's energy balance? What other bodily processes that are interconnected will be effected? Then you get into all the effects of going back into a gravity well. Would you come off the drugs (and thus require a washout period before you go downside, and a ramp-up period before you could go topside again)?

     Spin and gravity is an engineering headache, but a solvable one. Pharmacologically altering the body to prevent the loss of muscle and bone mass that the body seems surplus to requirements has all kinds of unknowns, off target-effects and unintended consequences. You're going to put people at severe risk for medical complications, some of which could be lifelong or even lethal.”

---

This is a compelling case that it is not possible to treat the effects of zero-G medically.  However, if for story reasons a workaround is needed, medical treatment is no less plausible than many devices used even in relatively hard Sci-Fi.

The task of designing spacecraft for a sci-fi setting is complicated by the need to find out all the things that need to be included, and get numbers for them.  The author has created a spreadsheet to automate this task, including an editable sheet of constants to allow the user to customize it to his needs.  The numbers there are the author’s best guess for Mid-PMF settings, but too complicated to duplicate here.

Rick Robinson’s general rule is that spacecraft will (in the sort of setting examined here) become broadly comparable to jetliners in cost, at about $1 million/ton in current dollars.  This is probably fairly accurate for civilian vessels, at least to a factor of 3 or so.  Warships are likely to be more expensive, as most of the components that separate warships from civilian ships are very expensive for their mass.  In aircraft terms, an F-16 is approximately $2 million/ton, as is the F-15, while the F/A-18E/F Super Hornet is closer to $4 million/ton.  This is certainly a better approximation than the difference between warships and cargo ships, as spacecraft and aircraft both have relatively expensive structures and engines, unlike naval vessels, where by far the most expensive component of a warship is its electronics.  For example, the ships of the Arleigh Burke-class of destroyers seem to be averaging between $150,000 and $250,000/ton, while various cargo ships seem to hover between $1000 and $5000/ton.

As mentioned in Section 5, some have suggested that the drive would be modular, with the front end of the ship (containing weapons, crew, cargo, and the like) built separately and attached for various missions.  This is somewhat plausible in a commercial context, but has serious problems in a military one.  However, the idea of buying a separate drive and payload and mating them together is quite likely, and could see military and civilian vessels sharing drive types.  (This is not as strange as present experience would lead us to believe.  It was only during WWII that military aircraft clearly separated from civilian ones in terms of performance and technology.)  This simplifies design of spacecraft significantly, as one can first design the engine, and then build payloads around it.

One common problem during the discussion of spacecraft design is the rating of the spacecraft.  With other vehicles, we have fairly simple specifications, such as maximum speed, range, and payload capacity.  However, none of these strictly applies in space, and the fact that spacecraft are not limited by gravity and movement through a fluid medium makes specifying the equivalents rather difficult.  Acceleration and delta-V obviously depend on the masses of the various components, which can be changed far more readily than on terrestrial vehicles, and cargo capacity is limited only by how long you’re willing to take to get where you’re going.  A replacement might be a series of standard trajectories, and the payload a craft can carry on them.  This works well if all of the spacecraft being rated are generally similar in terms of performance, and take similar trajectories in reality.  However, it does not work as well in a scenario where different types of ships take wildly different trajectories with different amounts of cargo.  In that scenario, ships might be rated by the minimum time for certain transfers (Earth-Mars at optimum, for instance) with a specified payload, either a fixed percentage of dry mass, or a series of specified masses for various sizes of ships.  This allows a comparison between ships of different classes, but within a class (liner, bulk cargo, etc.) the first method would probably be preferred.

A related problem is the selection of an appropriate delta-V during preliminary design.  In some cases, this is relatively easy, such as when a spacecraft is intended to use Hohmann or Hohmann-like trajectories, as numbers for such are easily available.  But such numbers are inadequate for a warship, or for any ship that operates in a much higher delta-V band, and unless the vessel has so much delta-V and such high acceleration that Brachistochrone approximations become accurate (and even then, if the vessel is not using a reactionless drive, the loss of remass can throw such numbers off significantly, unless much more complicated methods are used, numerical or otherwise).  The author has attempted to fill this gap by creating a series of tables of delta-Vs and transit times between various bodies, with the tables giving the percentage of the time that a vessel leaving one body can reach another within a specified amount of time with a given amount of delta-V.  The tables can be found at the end of this section.  Table 9 covers Earth-Mars transits, while Table 10 describes Earth-Jupiter transits.

The tables are generated in MATLAB by solving Lambert’s problem for a large number of departure days and transit times, and calculating the delta-V to go from stationary relative to the departure planet to stationary relative to the destination planet.  This involved assuming that there was a single instant delta-V burn at each end, which is a good approximation if the burn time is short compared to the transit time, as it would be for chemical or most fission-thermal rockets.  For systems which burn a significant amount of the time, this approximation is not as good, and the tables should only be used as a general guide to the required delta-V.  

Each table is the composite of 16 different tables generated with different starting geometry, and with each table containing data from at least one synodic period.  Note that this was all done in a sun-centric system, and that the delta-V necessary to deal with either planet’s gravity well was not included.  This will add some extra delta-V, the necessary amount shrinking in absolute terms as the overall delta-V is increased due to the Oberth effect.  The decision to not include escape and capture delta-V was made because to do otherwise would have involved specifying reference orbits to escape from and capture to, and would have added significant complexity to the program at a minimal gain in utility for most users.  

One thing that is apparent from these tables is the degree to which Jupiter missions are more hit-or-miss than Mars missions.  For Jupiter transfers, 84% of the options are either going to be viable all of the time, or not going to be viable at all.  For Mars, the equivalent value is only 56%.  Some of this is due to the much larger and more variable time increments used in the Jupiter calculations, but much of it is due to the fact that the geometry changes significantly less between Earth and Jupiter than it does between Earth and Mars.

It should also be noted that these tables are an attempt to find an average over all possible relative positions of the two bodies.  For the design of an actual spacecraft, analysis would instead start with modeling of geometries over the projected life of the spacecraft.  The approximations given here are reasonably close for theoretical use, but should not be used to plan actual space missions.  

Heat management is a vital part of the design and operation of a space vessel, particularly a warcraft.  Section 3 mentioned some of the issues with regards to stealth, but a more comprehensive analysis is necessary.  There are two options for dealing with waste heat in battle: radiators and heat sinks.  If the waste heat is not dealt with, it would rapidly fry the ship and crew.  

All space vessels will need radiators to disperse the heat they produce as part of normal operations.  If using an electric drive, power (and therefore waste heat) production will be no higher in battle then during cruise.  This would allow the standard radiators to be used indefinitely during battle without requiring additional cooling systems.  The problem with radiators is that they are relatively large and vulnerable to damage.  The best solution is to keep them edge-on to the enemy, and probably armor the front edge.  The problem with this solution is that the vessel is constrained in maneuver, and can only face one (or possibly two) enemy forces at once without exposing the radiator.  If the techlevel is high enough to make maneuver in combat a viable proposition, then radiators are of dubious utility in combat.  On the other hand, the traditional laserstar battle suits radiators quite well.  The faceplate and the forward edge of the radiators are always pointed at the enemy, and almost all maneuvers are made side-to-side to dodge kinetics.  The only problem is vulnerability to a direct kinetic hit.  If a projectile were to arrive precisely edge-on, it could tear the entire radiator in two.  Bending the radiator slightly would eliminate this vulnerability, but would also increase armor requirements.  However, even a bent radiator would still have issues with grazing impacts.  A projectile coming in very close to parallel with the radiator’s surface would tend to tear a long hole in it, as opposed to the small hole left by a projectile traveling perpendicular to the surface.  However, the low-incidence projectile would have to penetrate much more material, so kinetics designed for such attacks would naturally have more mass or fewer pieces of shrapnel than one designed for normal attacks.

Heat sinks avoid the vulnerability to damage of radiators, but have a drawback of their own.  By their very nature, they have a limited heat capacity, which places a limit on how much power a ship can produce during an engagement, and thus on the duration of an engagement.  If the heat sinks fill up, the ship would begin to fry unless the radiators were extended immediately.  In the game Attack Vector: Tactical, extending the radiators is used to signal surrender.  Obviously, the heat clock is a major disadvantage, but it is necessary when the vessel expects to expose several aspects to the enemy.  

One topic that briefly needs to be addressed is electric propulsion. In discussions, VASIMR-type engines are usually considered the baseline.  However, Dr. Joshua Rovey of Missouri S&T told the author that Hall Effect thrusters today are capable of the sort of performance that VASIMR is currently promising after development is finished.  VASIMR is apparently getting attention due to good marketing people.  

Another topic that deserves discussion is the effect of nuclear power on spacecraft design.  For large warcraft, nuclear power, both for propulsion and for electricity is a must-have.  Even if the design of solar panels advances to the point at which they become a viable alternative for providing electrical propulsion power in large civilian spacecraft, there are several major drawbacks for military service.  The largest is that solar panels only work when facing the sun, unlike radiators, which work best when not facing the sun.  The distinction between the two is important, as it is nearly always possible to find an orientation which keeps the radiator edge-on to the enemy and still operating efficiently, while a solar panel must be pointed in a single direction, potentially exposing it to hostile fire.  A solar panel is particularly vulnerable to laser fire, as it is by nature an optical device.  While hard numbers on this are surprisingly difficult to find, it appears that damage will probably occur to photovoltaics when exposed to intensities of around 300 KJ/m², for short pulses (<10^-4 seconds), with threshold requirements increasing from there as the pulse length increases.  For a CW laser (>1 second), the power flux required for damage is approximately 10 MW/m².  Photovoltaics can also be attacked using small particles such as sand, as described in Section 7 for use against lasers.  While a full analysis of the potential damage is beyond the scope of this section, it appears that sand would be a reasonably effective means of attacking photovoltaics, particularly given the large area involved.  The size of a solar array also complicates maneuvering the panels edge-on to the incoming particles, and could potentially raise structural concerns.

Radiators, on the other hand, are more resistant to damage.  Firing lasers at them will only decrease the thermal efficiency of the reactor slightly, as the radiator is designed to disperse heat.  Particle clouds that are designed for surface effects would be ineffective against a properly-designed radiator, or at very best reduce the emissivity by a small amount.  Small pieces of shrapnel designed to pierce the radiator entirely would be the best means of attack (described above), as fully armoring a radiator is likely to be impossible because of the mass requirements.  

However, it could be argued that this ignores the vulnerability of the reactor itself to damage.  While in Hollywood, “They’ve hit the reactor!” is usually followed by a massive explosion, that is not the case in reality.  First, the reactor is a very small target, usually shielded by the bulk of the ship, so it’s unlikely to be hit in the first place.  Second, nuclear reactors simply do not turn into bombs under any circumstances, and particularly not random damage to the core.  The few cases in history in which a reactor has gone prompt critical (SL-1 and Chernobyl being the best-known) were caused by poor procedure, and are vanishingly unlikely to happen due to random damage.  

That said, it still seems a potentially bad idea to put all of one’s eggs in a single basket.  Solar panels are highly redundant, but the reactor could still be put out of action with a single hit.  The response to this is fairly simple.  First, there are reactor designs using heat pipes that have sufficient redundancy to continue operation even if the reactor core itself is hit.  The specific heat pipe will be put off line, but if the design has 150, that’s not a great worry.  Second, the reactor and associated gear (power converters and such) are buried deep in the ship, where they will be difficult to get at, and the power converters can be duplicated for redundancy.

One last concern is the ejection of reactor core material after a hit, and the potential for said material to irradiate the crew. This is also probably minor, as the crew is still being shot at, and the spacecraft will have some shielding against both background radiation and nuclear weapons. (Thanks to Dr. Jeffrey King of the Colorado School of Mines for providing much of the material on space nuclear power and propulsion.)

A couple of other issues with nuclear power are relevant and of interest.  The first is the choice of remass in nuclear-thermal rockets.  While hydrogen is obviously the best possible choice (the reasons for this are outside the scope of this paper, but the details are easy to find), it is also hard to find in many places.  With other forms of remass, the NTR does not compete terribly well with chemical rockets, but it can theoretically use any form of remass available.  The biggest problem with alternative remasses is material limits.  With most proposed materials, oxidizing remasses will rapidly erode and destroy the engine.  The alternatives to avoid this are rhenium and iridium, which are both very expensive, explaining why they are not in use today.  However, both elements are common in asteroids, making them viable choices in a setting with large-scale space industry.

As discussed in Section 7, vibration is a serious issue for laser-armed spacecraft.  Any rotating part will produce vibrations, and minimizing these vibrations is of interest to the designer.  While there is undoubted a significant amount that could be done to reduce the vibrations produced by conventional machinery (the exact techniques are probably classified, as their primary application is in submarine silencing), it seems simpler to use systems with no moving parts, which should theoretically minimize both vibration and maintenance.  Heat pipes, as mentioned above, are an entirely passive means of moving heat around, both from a reactor to an energy converter (which could mean a turbine, a thermocouple, or any of the other wonderful things engineers can think of) and from the energy converter to the radiator.  There are also electromagnetic pumps for liquid metal which while not entirely passive, but will cut down on the vibration load.

There are even proposed systems of energy conversion which are reasonably efficient and involve no moving parts.  The best-known of these proposed systems is probably the Alkali Metal Thermal-to-Electric Converter (AMTEC), which has been extensively studied.  However, a recent effort by NASA to bring the technology into deployment failed, giving the technology a bad name.  There are some, however, who believe it still holds promise.

If systems like AMTEC are not available, the spacecraft will have to use conventional hat engines.  These are likely to use one of the two standard thermodynamic cycles, the Brayton cycle (gas turbine), and the Rankine cycle (steam turbine).  The primary difference between the two is that in the Brayton cycle, the working fluid remains a gas throughout, while in the Rankine cycle, it moves from liquid to gas and back again.  In theoretical design, radiators are normally sized assuming constant temperature throughout, which is true for most Rankine cycle systems (as the radiators are where the fluid condenses at a constant temperature) and produces the well-known result that radiator area is minimized when the radiator temperature is 75% of the generation temperature.  However, this is not true for Brayton cycle radiators.  There is no convenient mechanism to release the necessary heat at a constant temperature, so the radiator performs differently as the gas cools.  There is not a simple formula here, but an iterative procedure can be used to minimize radiator area for an ideal system (which is close enough for our purposes).  Use of this method does require some knowledge of the basics of gas turbine propulsion, but it is not terribly esoteric. (Thanks to Dr. David Riggins of Missouri S&T for presenting this material in class. Those who have had more experience in propulsion and fluid dynamics might recognize some simplifications of the explanatory material, and some nomenclature changes. This was intended to hold down the length of this section and clarify it without sacrificing accuracy of the results.)

An ideal gas turbine can be thought of as being made of 4 separate stages.  First, isentropic compression, which means that there is no heat transfer and all energy put into the system by the compressor, is used to compress the gas instead of heating it.  Second, isobaric (constant-pressure) heat addition, which occurs in the reactor.  Third, isentropic expansion through a turbine, which outputs mechanical work (the goal of this whole process).  Lastly, isobaric heat rejection, through the radiator, which returns the working fluid to the condition it was at before entering the compressor.  The compressor and turbine are defined by their pressure ratios, written as π_c and π_t respectively.  The pressure ratio is the pressure of the fluid after the component divided by the pressure ahead of the component.

For this method, values for C_p, γ, η_c, η_t, and T₃ must be selected.  C_p is the constant-pressure specific heat capacity of the fluid, while γ is the ratio of specific heats.  Definitions and values for various fluids can be found online.  η_c and η_t are the efficiencies for the compressor and turbine respectively.  Values between 0.8 and 0.9 are probably reasonable.  T₃ is the temperature at the outlet of the heat addition stage, and is normally set by the design of the reactor itself.  Representative values might be 1600-1700 K for a conventional nuclear reactor, although higher values are possible.  (All temperature values throughout should be in Kelvin, not Celsius.)  

The first step is to select a value for π_c, with anything from 2 to 10 being plausible.  Because it is a closed system, π_t will be equal to 1/ π_c.  Once this is known, it is possible to calculate T₄ (temperature downstream of the turbine) using .

At this point, a value for T₁ must also be selected.  This is the temperature at the entrance to the compressor.  Using this, the value for T₂, the temperature at the compressor exit, can be calculated using .  This allows the overall efficiency of the power-generation system, &eta (work output/heat input), to be calculated using .  All of this information can then be used to find the radiator area per unit work output (A’, m²/W) with where σ is the Stefan-Boltzmann constant (5.670373×10^-8) and ε is the emissivity of the radiator (0.9-1.0). Once you have this value, select a different value of T₁, and repeat the rest of the paragraph.  When a minimum has been found, select a different value for π_c and repeat the entire procedure until a global minimum has been found.  It would probably be a good idea to use a spreadsheet to automate this.

Table 9: Earth-Mars Transits

ΔV km/s	Maximum Transit Time (Must Arrive No Later Than, Days)
0	30	45	60	90	120	150	180	210	240	270	300	330	360
10	0	0	0	0	0.4487	1.9151	6.4984	12.019	16.394	20.353	24.215	28.229	32.3
20	0	0	0	2.0593	9.2949	16.202	23.165	32.821	41.09	48.125	54.311	59.856	65.337
30	0	0	1.0497	9.2548	17.492	27.444	43.646	57.051	66.458	74.976	82.428	89.111	95.232
40	0	0.8333	5.625	13.918	26.034	48.325	66.667	77.019	86.747	95.016	98.854	99.952	100
50	0.072	4.2468	8.5978	18.454	39.03	69.872	81.202	90.601	95.938	98.646	99.944	100	100
60	0.617	6.2821	10.929	24.615	67.067	80.545	89.303	95.409	98.934	100	100	100	100
75	3.6458	8.8061	13.99	44.896	79.191	88.678	95.28	98.702	99.984	100	100	100	100
100	6.3141	12.22	21.811	81.466	91.603	96.851	100	100	100	100	100	100	100
125	8.5176	16.763	50.529	92.388	100	100	100	100	100	100	100	100	100
150	10.401	23.301	86.787	100	100	100	100	100	100	100	100	100	100
175	12.724	53.886	100	100	100	100	100	100	100	100	100	100	100
200	16.226	90.088	100	100	100	100	100	100	100	100	100	100	100
250	26.963	100	100	100	100	100	100	100	100	100	100	100	100
275	58.389	100	100	100	100	100	100	100	100	100	100	100	100
300	94.255	100	100	100	100	100	100	100	100	100	100	100	100

Table 10: Earth-Jupiter Transits

ΔV km/s	Maximum Transit Time (Must Arrive No Later Than, Days)
0	30	40	50	60	75	100	125	150	225	300	450	600	750	900	1050	1200
15	0	0	0	0	0	0	0	0	0	0	0	0	0	2.949	15.32	28.99
20	0	0	0	0	0	0	0	0	0	0	0	7.055	46.48	86.71	99.61	100
25	0	0	0	0	0	0	0	0	0	0	0.593	35.77	76.52	98.68	100	100
30	0	0	0	0	0	0	0	0	0	0	14.72	54.85	92.54	99.82	100	100
40	0	0	0	0	0	0	0	0	0	0.666	37.92	79.23	99.07	100	100	100
50	0	0	0	0	0	0	0	0	0	13.03	53.12	90.34	99.90	100	100	100
60	0	0	0	0	0	0	0	0	4.351	22.86	68.86	96.05	100	100	100	100
75	0	0	0	0	0	0	0	0	15.02	45.55	95.53	100	100	100	100	100
100	0	0	0	0	0	0	0	6.385	52.61	93.81	100	100	100	100	100	100
125	0	0	0	0	0	0	6.559	17.55	96.00	100	100	100	100	100	100	100
150	0	0	0	0	0	4.020	16.43	69.32	100	100	100	100	100	100	100	100
175	0	0	0	0	0	10.60	60.9	100	100	100	100	100	100	100	100	100
200	0	0	0	0	2.846	22.82	100	100	100	100	100	100	100	100	100	100
250	0	0	0	2.238	15.07	100	100	100	100	100	100	100	100	100	100	100
275	0	0	0	6.236	39.69	100	100	100	100	100	100	100	100	100	100	100
300	0	0	1.901	11.54	91.61	100	100	100	100	100	100	100	100	100	100	100
350	0	0	9.396	49.64	100	100	100	100	100	100	100	100	100	100	100	100
400	0	4.240	22.52	100	100	100	100	100	100	100	100	100	100	100	100	100
500	1.429	22.66	100	100	100	100	100	100	100	100	100	100	100	100	100	100

• 

• 

Well, I was going to post the second part of The Shipping Trade today, except that writing it didn’t happen because of day job, and so forth. Then, I thought I might post a sketch of the ship involved, just to give y’all an idea of what you’ll be looking at, but then that would require me to go out and hire a scanner. That, and I made said sketch, and then looked at it, and then concluded that I couldn’t possibly inflict such a terrible picture on my readers…

So permit me, please, instead to sketch a verbal picture for you of the

---

CMS Greed and Mass-Energy

To start with, Greed and Mass-Energy is atypically large for a free trader; in those leagues, which principally deal in small, high-value-to-mass/volume cargoes, lugging around 40,000 tons displacement of cargo is huge. (It’s still not in the major freight line league, though; those guys can use freighters that are million ton-displacement behemoths.) Thus, the shipcorp that owns her (it’s essentially a syndicate of officers, crew, and former crew, with executive power vested in the captain-owner) is pretty prosperous to be able to cover her running costs. Dealing in brokered cargo actually isn’t her main business – she specializes in contracts like the RCS-assembly charter from Kerbol to Kythera she just left, but an empty hold is a hole that drinks money, so you take the cargo when you can get it.

Also, obviously, at a size like that, she’s not streamlined, or built to land planetside (gravity wells being acutely expensive); and is even rather more massy than anything that most stations like to have dock directly to them. Her cargo’s generally ferried to station, or upwell and downwell, by local lighters at each end of the trip. Rather, she’s built very much in the classic mode; a long, relatively thin, open-frame truss structure. Attached to that, going from fore to aft, we find these different sections of the ship:

Right at the bow, sitting on the end of the main truss, is the command capsule, an ellipsoid slightly stretched along the ship’s main axis, relatively tiny compared to the rest of the ship, and containing, for starters, the bridge and associated avionics systems. (The bridge is actually buried in the center of the capsule, for its protection; it’s displaced off to the front end of the ship, however, because the command capsule is also where the primary sensors are housed to keep them out of the way of cargo, fuel, and drive radiation, and this positioning cuts down on sensor lag. It’s still pretty safe; it’s not like anyone’s going to be shooting at them.) The first of the other two notable features it houses is docks and locks, right for’ard on the axis where it’s easiest to match thrust and spin, which usually houses a couple of cutters used for taking the crew ashore and for occasional maintenance, and a skimmer for in-field refueling. (The fuel itself doesn’t pass through here – the skimmer docks aft to offload what it scoops. No fuel for’ard of the support plate, that’s the general rule.) The second, aft by the truss, is the robot hotel for all the little space-rated utility spiders you may see now and them crawling about the structure doing maintenance, thus saving the engineering department any need to get suited up and go outside for routine work, although they still may need to do so from time to time.

Just aft of that, accommodations and secondary systems are housed in a toroidal gravity wheel. This is actually a very unusual design feature in an Imperial ship-class; just about everyone and especially the spacer-clades are genetically adapted to microgravity, and the spacer-clades prefer it, as a rule; but the Cheneos-class architects originally designed her class for near-frontier work, and included this for occasional passenger service. Greed and Mass-Energy only rarely carries passengers, so they keep it geared all the way down, producing only a tenth of a standard gravity, which doesn’t offend the spacer-clades all that much. There’s a second, smaller wheel rotating inside it to null out the gyroscopic effects; it’s used to house some other equipment that likes a little gravity, but for the most part, this one’s just a countermass.

(The wheel does, however, provide enough gravity to let the CELSS Manager run a pretty decent microbrewery in the spare volume, and perhaps more importantly, provides a place where you can drink it off-shift without suffering from a nasty case of the zero-g bloat. [Remember, folks, bubbles don’t rise in microgravity!] And apart from crew morale, having decent beer makes for good PR when traders meet.)

These areas, incidentally, are one of the few places on board where the really high-tech ontotechnological stuff makes an appearance, in the form of inertial damping. The people who built her liked microgravity, and weren’t all that keen on losing that while under thrust, especially since she was built to fly brachistochrones or near-brachistochrones (bulk tankers and ore freighters, etc., are usually built to fly economic minimum-delta/Hohmann transfers; no-one else wants to wait that long for their cargo) and so would be spending most of her time under thrust. The job of the inertial dampers is to apply the thrust of the drives evenly across the entire area’s structure and everything in it, thus ensuring that no-one actually feels any acceleration, and the lovely microgravity environment is preserved. (It also avoids having to come up with some wretchedly complicated gimbal arrangement for the already wretchedly complicated seals-and-bearings for the gravity wheel, no longer having to do which is something that made architects particularly grateful for this innovation.)

Behind this, the cargo. ‘Way back along the truss there is a very large, solid plate, the support plate. The cargo containers are simply stacked “atop” – by which we mean for’ard – of it, in six big blocks arranged around the axis with sixfold symmetry (this arrangement being a reasonable compromise between use-of-volume and convenient straight lines), and are designed to lock to the plate, the truss, and each other to form a solid interlocked structure. There’s no hold or other walls around the cargo; the containers are themselves spacetight when they need to be, and so lighters can just drop them into place and pick them up freely while in port.

The breakbulk cargo, on the other hand, is messy. It has to be podded up individually when not spacetight, and then individually lashed down and made secure atop the cargo container stacks. This annoys the cargomaster, which is why breakbulk is unpopular these days despite the fact that breakbulk shippers usually pay a premium in exchange for you having to do this (the “lash comp”). Actually, what really annoys the cargomaster is that she can punch a button and have the ship automatically query the v-tags on the container cargo for its mass stats, and so forth, whereas for breakbulk she’s got to recall her Academy training, dig out the spreadsheets, and work out the corrections to the center-of-mass-and-moment-of-inertia chart by hand. Well, still by computer, but you know what I mean.

Aft of the support plate, still in sixfold symmetry, you have the bunkerage – fuel tanks, stacked three deep in multiple rows, all filled with slush deuterium, running right to the stern, where they surround the cylindrical shroud of the mostly-unpressurized engineering hull (you can take a crawlway right back along the truss to the small, pressurized maneuvering room back this far, should you need to examine the drives close-up in flight, but the actual machinery space isn’t), which contains the interlinked systems of the main power reactors and the fusion torches themselves, strapped to the aftmost extent of the main truss.

And there are lots of fuel tanks. Even though said fusion torches are miracles of a mature nuclear technology, capable of achieving near-theoretical efficiencies and outputs and delta-v per unit fuel that routinely makes naval architects from less advanced civilizations throw down their slide rules in despair and weep into their terrible coffee-equivalents, the one unchangeable rule of space travel is that your mass ratio is always much, much less favorable than you might want it to be.

Good thing deuterium’s so cheap, isn’t it?

…and most prominently of all from a distance – dominating the entire view of the ship from a distance, by area as well as by temperature – sweeping out from among the fuel tanks (although comfortably retracted to sit alongside them, leaving approximately a sixth of their radiative area useful, while idling in dock – the vast panels and pipework of the heat radiators. Because the other one unchangeable rule of space travel is that you always have waste heat, too damn much waste heat, and you’ve got to get rid of it somehow. Especially once you fire up those fusion torches. (The radiators, however, unlike the rest of the ship, have only fourfold symmetry – so that they can be perpendicular to each other when unfolded, because there’s very little point in radiating heat right back at your own radiators.)

      Now, to regress to the region of solar system transportation. I'm going to confine my discussion almost completely to gaseous fission rockets. The reason is not that I have decided that they are the things to be used rather than Orion or electrical rockets. The reason is simply that I have done more thinking about them. I think that I can see how to combine gaseous fission engines with advanced vehicles more efficiently than is possible with the other engine types. I believe that not enough thought has been given to the engine/vehicle interaction. There was a statement in the letter of invitation to this symposium to the effect that everybody knows what to do with 1800 seconds specific impulse. I happen to disagree. I don't think anybody knows what to do with 1800 seconds specific impulse, and I think we wouldn't know what to do with a good space engine if it walked up and bit us. I will try to prove those opinions as I go. I've used the "we" advisedly because I don't think I know either. Rather than getting into a big mish-mash by attempting to cover all various forms of propulsion, I will just stick with some mythical gaseous fission rockets. Presumably in the next 3 days we will discuss which engines you really should do this with, if you should do it at all.

• 

     Figure 6 is a presentation of operating cost in dollars per pound of payload versus total velocity capability for single-stage vehicles with chemical, nuclear solid-core, and two different kinds of gaseous fission rocket engines. These curves were calculated four years ago when I was at Douglas for a paper by myself, Bill Mathiesen and Bob Trapp that was given at the I.A.F. Congress in Stockholm. We thought this was pretty interesting, but other than shocking an occasional person here and there, not very much has happened as a result. The thing that has been interesting to me is the fact that here was a clear indication that one could get out to extremely high velocities for a very low transportation cost. Velocities so high that you could open up the whole solar system for exploration with reasonable costs. Yet almost everyone believes it is extremely hard to do a little bit of space flight down in the low velocity region where we talk about just barely going to the moon.

     Figure 6 also is an interesting indication of the fact that you shouldn't drive a rocket faster than it wants to be driven. This is something that apparently a lot of people are forgetting. Theoretically, a rocket can go up to any velocity, not counting Einstein. But the way to make an inefficient rocket go to very high velocity is to stage it, and stage it, and stage it. One carries fuel to carry the fuel to carry the fuel that's going to be used later. Once the weight starts to pyramid, it's a logarithmic function and it just plain gets ridiculous in a hurry. So one has to be very careful about taking something like a solid-core nuclear rocket and deciding to perform missions at 200,000 fps. You can stack tip all that equipment if yqu want to do it, but it's a horrendous thing.

     A question that has bothered me quite a bit is why more attention was not paid to these curves. Consequently, I'm going to break down some of the assumptions used in these curves, then build new curves back up with this year's assumptions. Perhaps I can make the story more believable.

     Now, there are two big ringers in the curves of Figure 6. One is the obvious one. At that stage of the game, nobody had the foggiest idea of how to build a gaseous fission engine at all, let alone one with 5,000 to 20,000 seconds specific impulse. That, right off the bat, caused everybody to throw up their hands and forget it. The second ringer is that we used transport airplane operating cost assumptions. To put it mildly, we used recovery and reuse assumptions which were not the standard thing in rocket work. I'm going to examine both of these assumptions in today's light.

• 

     Since so little was known of gaseous fission engines four years ago, the previous study assumed a thrust/weight ratio of 30 independent of specific impulse. This was recognized to be a very sporty assumption since, even if the containment problem coUld be solved, the achievement of specific impulses beyond about 3,000 seconds requires the use of a radiator to reject excess heat which cannot be handled by the thermal capacity of the propellant utilized (excess heat above what can be handled by open-cycle cooling). Although it was originally felt that such radiators would represent an intolerable decrease of thrust/weight ratio, it has since been pointed out that this is not true if high temperature radiators are used. Figure 7 shows a current estimate of the variation of the thrust/weight ratio with I_sp achievable for a gaseous nuclear rocket system with radiator using as a basis an assumed thrust/weight ratio of 20, at an I_sp of 2,500 seconds. The values fall off substantially at high specific impulses compared to the assumptions of 4 years ago, but are still greater than one to beyond 10,000 seconds specific impulse.

     It should be pointed out that the use of water, ammonia, methane, or other non-hydrogen working fluids should be seriously considered in gaseous fission engines from the start. Not only are better ship designs permissible due to small tankage sizes and ease of propellant storability, but the use of a higher density propellant might well ease the fuel containment problem if a vortex system is used. If so, it could result in smaller, lighter engines. It might also result in earlier development programs if the ability to prove adequate containment occurred at an earlier time.

(ed note: In thermal rockets, hydrogen produces the best specific impulse. But on the minus side: hydrogen tanks have to be huge because of low density, and it is a pain in the ass to store because is is cryogentic and needs lots of refrigeration)

     An interesting result in Figure 7 is that the value of thrust/weight ratio as the specific impulse approaches 10,000 seconds is independent of propellant used. This is because the reduced propellant flow at such a high specific impulse results in such small thermal capacity in the incoming fuel that the engine must be almost completely cooled by the radiator system. The propellant to fuel burned ratio required for a given specific impulse is independent of propellant used. Hence, the engine uses the same amount of energy to generate a given specific impulse, the same fraction of energy must be rejected by the radiator, and the radiator area is unaffected by the type of propellant used.

     It seems clear that an engine design cooled by radiator alone should be investigated. Such an engine might be easier to develop since a major interaction between propellant and cooling system would be severed. Furthermore, such an engine could more easily use a variety of propellants. This could be very helpful in early planetary exploring.

     A limitation on specific impulse of 10,000 seconds has been shown tentatively in Figure 7, assuming that the engine would be of the type which transfers heat from the fission plasma to the propellant by radiation. This is due to an unfortunate tendency of the propellants examined to date. Although adequately opaque to absorb the radiant energy at medium-high temperatures, they apparently become transparent at very high temperatures. At the moment seeding the flow, which is very effective at low temperatures, does not look promising at high temperatures.

(ed note: "seeding the flow" means putting tungsten dust in the propellant. Otherwise the heat radiated by the nuclear reaction fails to heat up the transparent propellant, and instead the rocket engine is destroyed)

     One other point of interest in connection with the thrust/weight ratios of gaseous fission engines is the power conversion weight thus achieved. Electrical propulsion enthusiasts feel extremely optimistic when power conversion weights of the order of 10 pounds per kilowatt are mentioned. A gaseous fission engine of 2,500 seconds I_sp and T/W of 20 achieves about one-thousandth of a pound per kilowatt. In other words, gaseous fission engines are almost certain to be 10,000 times better than electrical rockets in power conversion weights. The fabulous effect of this number on spaceship design must be understood if anyone expects to make rational development decisions on future propulsion systems.

     I can defend the 1963 curves of Figure 7 today. Not very well, but at least I can begin to defend them. Four years ago, the 1960 assumption was nothing that could be defended at all. However, I have always liked the calculations we made then. It influenced me in feeling strongly that, at least theoretically, there was a great deal more which could be done with nuclear rockets than anyone realized, far more than just trivial improvements in our current systems.

     As part of the process of understanding spaceship operating costs, it is instructive to consider first only the fuel and propellant cost. This is true because this cost represents the minimum achievable. It is important to understand the mechanics of achieving a low fuel and propellant cost, particularly when truly reusable ships are used. In transport aircraft practice, the amount of reuse is so high that initial airframe costs are only a small fraction of the operating cost, and fuel costs represent about one-half the total. Thus, we shall examine fuel costs for their basic limitations on performance, and then see how closely these limits can be approached with reusable ships.

• 

     Fuel and propellant costs as a function of total velocity increment for chemical, solid core nuclear, and gaseous nuclear rockets are shown in Figure 8. Compared to the assumptions of Figure 6, the specific impulse of the high energy chemical has been increased to represent a modern, high-pressure system: the solid core nuclear has been decreased in view of current development difficulties: and a number of different propellants and degrees of containment are shown for gaseous fission engines. All curves are for single-stage ships with structural assumptions more conservative than those of Figure 6 and each specifically sized for the velocity shown.

     It is evident that, on this basis alone, a gaseous fission engine without radiators and with separation ratio of 10^-3 is not significantly better than a solid core engine. Gaseous engines with better containment would be much better. It is also evident that gaseous engines with space radiators, but with specific impulse limited to 10,000 seconds, can drive ships up to about one-half million feet per second and still maintain reasonable fuel cost. The attainment of a fuel separation ratio of 10^-4 is almost as effective as perfect fuel containment.

     The optimum fuel cost curves for gaseous fission engines with radiators were obtained by determining the optimum specific impulse for each velocity and separation ratio. This is necessary since too low a specific impulse will result in excessive propellant cost while too high a specific impulse will result in excessive fuel cost. The optimum specific impulse is much higher than 10,000 seconds for all velocities beyond a few hundred thousand feet per second. Hence, these curves represent a future capability presently unattainable due to the propellant transparency problem at high temperatures previously mentioned. If it were not for this, gaseous fission ships could be driven to almost one million feet per second before fuel costs became a limitation.

     Under certain circumstances, a great deal can be learned about spaceship design without a detailed knowledge of the missions to be performed. In recent years, there has been a tendency to become so detail-mission-oriented that ship design is not even attempted until the exact mission is clearly understood. This may be a valid procedure when only one mission is in sight, although even then the inevitable lack of versatility usually leads to needless redesign much earlier than anticipated. In transport airplane design, the basic design procedure usually centers around the calculation of airplane operating characteristics as a function of range. The maximum range required comes from a knowledge of the total mission complex, but the airplane design is refined primarily by using general curves as a function of range, rather than by a detailed series of specific mission analyses.

     When considering total solar system transportation as we are, it is. clear that we face a variety of missions. It is also very unclear as to which of these will be paramount. One way of approaching the problem is to present the characteristics of the vehicle as a function of total velocity increment which the ship can achieve. This is exactly analogous to the use of range in aircraft design practice. In this way, an understanding of the ship's basic ability to deliver payload to a certain speed economically can be rather easily understood. The complex mission analyses, then, can be made to reflect the maximum design velocity increment required.

     This approach was actually used in the 1960 study, and Figure 6 represents one of the results. Figure 6, however, contains assumptions as to degree of reuse achieved by the vehicle which, although consistent with transport aircraft practice, may not apply to space transportation. At least, if they do, their application must be better documented.

     The 1960 study assumed a large number of reuses per vehicle, somewhat analogous to the number of times a transport airplane is reused. A transport aircraft is actually utilized about 50 percent of the time, and average flight durations are less than 4 hours. It is clear, therefore, that such vehicles are used over one thousand times per year. However, space travel durations are much longer, and it is obvious that the interaction between travel duration and number of reuses must be considered.

     For the lunar mission, it is clear that large numbers of reuses are feasible. Typically, 100 flights per year (50 each way) can be envisioned on the basis of 2-day travel times, one day turn around time at each terminal, with Sundays and 2 weeks off for vacation. Over a 10-year ship lifetime, 1,000 uses will be achieved.

     One can get a feeling for the number of interplanetary uses by assuming a certain ship total life. Typically, transport aircraft are designed for 40,000 hours (4.6 years) total life. On the basis of slightly less than 50 percent utilization, such a vehicle would last for 10 years. They always last much longer, but the amortization time of the airframe is usually about 40,000 hours, since new equipment always becomes available in even shorter time.

     Selecting a suitable lifetime for a spaceship presents a considerable technical dilemma. One viewpoint would simply take 10 years as above. An even shorter lifetime might technically be justified due to the severe aerodynamic environments associated with atmospheric entries, and the generally unknown operational environment of space. This type of assumption has become standard in this country recently. If you don't understand the problem, assume it's horrible.

     It may well be, however, that spaceships will last much longer than transport aircraft. The transport has its main propulsion system operating continually during flight, and is also continually facing the temperatures and gust loads within our atmosphere. The question is whether spaceship operating life should be determined by the total time of operation, or only by the times during which the main engines operate and/or it is within an atmosphere. In other words, is a spaceship coasting between planets actually operating in the aircraft transport sense, or is it merely parked in space, breathing quietly, waiting for its next mission.

     One can make an excellent case for the latter point of view in terms of the general environment that the ship faces, either from space or its own propulsion systems, while coasting. The ship would have to be on interplanetary runs for several centuries in order to build up 40,000 hours of engine and atmospheric operation. It is, however, bound to be replaced by better equipment within a few decades. As a base for calculations, this report assumes 25 years ship useful lifetime.

• 

     The variation of various weights as a function of velocity is shown in Figure 9 for both specific impulse limited to 10,000 seconds and for the optimum specific impulse. These curves are for ships designed for 20 percent payload, then operated at lower velocities by off-loading propellant and at higher velocities by off-loading payload. Thus, these curves represent a penalty for using a single ship for multiple missions, just as in other forms of transportation.

• 

     By using suitable planetary travel time data, which is not yet easy to come by, the weight data of Figure 9, the same fuel plus propellant cost assumptions as in Figure 8, and assuming a vehicle cost of $100 per pound, the curves of Figure 10 were obtained. One hundred dollars per pound is the currently estimated cost of a supersonic transport. These curves show fuel cost plus amortized airframe cost as a function of design velocity increment for the missions selected.

     The lowest curves on Figure 10 are fuel cost only. Comparing them with the other curves show that for operations as far out as the planet Saturn, the structural costs are comparable to fuel costs. Further improvements in convenience of operation can be achieved with engines not limited to 10,000 seconds specific impulse. In that case, velocity increments beyond a half-million feet per second are economically reasonable.

• 

     The average travel time between planets corresponding to the velocities of Figure 10 are shown in Figure 11. These two Figures taken together give a better feel for solar system transportation than Figure 6 alone. With specific impulse limited to 10,000 seconds, the solar system as far as Jupiter is available with travel times not exceeding 4 months. Inner solar system travel times need not exceed 2 months. The advantage of optimum specific impulse becomes more evident at Saturn and beyond.

     The curves of Figures 10 and 11 apply for a given ship design velocity only if the ship can be refueled at each terminal. If it must carry its own fuel for the return journey, then it must operate at half the total velocity shown. Except for Pluto, refueling bases at the major planets are much more needed than at the minor ones, as can be seen by Figure 10. Refueling bases could be expected to be located on the surfaces of all the minor planets, although it may require some design effort in the case of Venus and Pluto.

     The major planets are a different situation. Their surfaces are extremely forbidding as far as we know, to the extent that we are not even sure they have solid surfaces. It makes sense in that case to establish bases on one of the satellites of each of the four major planets. The curves are drawn with that assumption. If we do decide to penetrate to the surface of these planets, then the velocity requirements for doing this when operating from one of the satellites is a reasonable number. Thus, bases on the larger planets' satellites not only greatly facilitate the convenience of transportation, but also present a reasonable base for surface exploration, if required.

     These particular curves are also calculated for the average flight times involved in year around operations between all planets. There are no launch window restrictions. If you're ever going to have a transportation system, you're going to have to be able to go when you want to. You cannot spend most of the time waiting. A rough averaging process between the best and worst times of the year was used in. an attempt to make this a realistic transportation assumption.

     A few words about perspective on these curves are in order. There is nothing magic here except a ridiculous willingness to plot curves wherever the data is leading, rather than stopping somewhere. Both better and worse situations may well occur. Even the case of specific impulse limited to 10,000 seconds requires gaseous fission engines with radiators, and most people today would rather agree to engines without radiators. In that case, the velocity increment achieved will be only about 25 percent of the curves shown. Furthermore, the economic penalty of, if necessary, ejecting a critical mass of fuel in the process of shutting down the engine has not been included. This will be on the order of $100,000 per shutdown ($778,160 in 2017 dollars).

     On the other hand, perfect containment might be achieved. We might design ships for each velocity increment, rather than use the single design assumed here. Furthermore, one can get a greater utilization of vehicles by the expedient of refueling the vehicles which go on deep space missions. This is preferable to multi-stage vehicles, since a fleet of ships used for refueling can also be used for other missions. No attempt will be made here to present detailed curves showing the effects of refueling. Cursory checks show that over 200,000 fps can be added for reasonable cost with only two refuelings.

     The greatest conservatism of all in Figures 10 and 11 is, of course, in the magnitude of the ordinate scale. Costs beyond $12 per pound have not been plotted so that the entire set of curves is about 100 to 1,000 times lower than virtually all space cost analyses to date. This must be clearly remembered as we discuss the performance of these ships.

     I can't resist making one more solar system point here. So far, only travel between Earth and the other planets has been discussed. There is also the question of travel between planets other than Earth. The use of bases in other parts of the solar system to aid in the exploration of the even more remote portions should be considered. In fact, such considerations might well dictate the strategic location of bases.

     At first thought, it would seem to be a good idea, for instance, to use a base on one of the farther planets, say Saturn, to permit further exploration of the more remote planets like Pluto. Although this is an intriguing thought, such deep bases will have only limited utility. The reason is the extremely long synodic periods which exist among the outer planets since they move so slowly around the Sun. In the worst case of all, the synodic period between Neptune and Pluto is slightly over 500 years. In addition to the long synodic period, the difference between travel at the optimum time of the year and the worst time of the year becomes more extreme the farther the planet is located from the Sun.

• 

     One way of illustrating this is shown in Figure 12, where the effects of basing on selected planets is shown for a constant ship velocity. It is true that a deep space base will be closer than Earth to the other deep space objects when in favorable position, but equally true that it will be much farther away during the worst conditions. Surprisingly enough, the base wants to be reasonably close to the Sun, once again emphasizing that the Sun is the center of the solar system. Although Mercury might be the best planetary base of all, the Earth is still sufficiently close to the Sun that it represents a pretty good compromise. Thus, the major space logistics support operations could, from a celestial mechanics viewpoint, be located efficiently on the Earth or its Moon. This is very convenient since the known industrial and research bases of the solar system also happen to be located in that vicinity.

     The slow movement of the outer planets leads to some interesting paradoxes. One would naturally assume that a base on Triton would be an excellent place from which to explore Pluto, since Neptune is at 30.09 A.U. from the Sun, while Pluto is 39.5 A.U. However, it turns out that Neptune at the moment is already leading Pluto around the Sun, and pulling away. In fact, in approximately 9 years, Neptune will be farther away from Pluto than Earth ever is. Furthermore, due to the long synodic period of Neptune and Pluto, that statement will be true for somewhat over the next 300 years. It would be nice to be sure that every other statement in this discussion will be true for that duration.

     I threw that in as a bit of tidbit. I' m gradually working back down to our more normal systems, and there's a point that I want to harp on further — that is, this whole question of structural reuse. As I indicated before, Figures 10 and 11 show costs that are less than $10 per pound throughout the entire solar system. Yet large and elaborate studies are made these days proving that it's going to take many hundred dollars per pound to go to the moon, no matter what we do with recovery, reuse, or anything else. It1 s quite clear that either I'm insane, or a lot of other people are, or we have to have an explanation. It was easier to give an explanation than to prove everyone else insane.

     Figure 6 was drawn with what I like to refer to as "transportation" type assumptions for operating cost. Maintenance costs, for instance, were taken from normal air transport practice. The vehicles were assumed to be as reusable as transport aircraft. This is not the fashion when calculating rocket operating costs today. Almost all of our rocket builders, including myself, have been doing nothing but "ammunition" work for a large number of years. It is axiomatic that, if you want to talk about recovery and reuse, you should not talk to an ammunition builder.

• 

     To illustrate what can happen with different classes of maintenance and reuse assumption, I will shift gears drasticaily and discuss merely placing objects on Earth orbit with chemical propulsion. In Figure 13, I assumed, arbitrarily, about $300.00 per pound of payload as typical of current day orbital transportation systems with no reuse at all. It so happens, however, that if you consider the actual price of high energy fuel needed for orbital velocities with advanced rockets, it is only on the order of $1.00 per pound of payload. Figure 13 is simply a plot of operating cost as a function of recovery reliability and refurbishment cost with these assumptions.

     It is quite fashionable, whenever over-all system analyses for recoverable space vehicles are performed, to assume that recovery reliabilities will be around 75 percent. After all, that is the recovery experience to date. Also, refurbishment costs around 25 percent are quite likely to be used. The shaded region brackets these assumptions, and is typical of a good, solid, rational ammunition type analysis. It is evident that after spending that much money on refurbishment between flights with that low a recovery reliability, an improvement of at most two in over-all cost performance is the best to be expected.

     Also shown on Figure 13 is what had already been achieved many decades ago in air transportation. This is what happens when you think like a transportation man. The recovery reliability is so close to 1.00 that you can't possibly see it on this scale. The same is true of the maintenance cost, which is on the order of 0.04 percent. If anyone here thinks that a DC-8 is less complicated than a Thor, just take a good look at the inside workings of a DC-8 some day. Wonder, then, at the fact that a few people turn it around, give it some fuel, pat it on the head, and it takes off again. This is what we should be trying for in future spaceships. There is an improvement of a factor of 100 over current operations to be made. I do not want you to get the impression that I am all for airplane designers. I think they, too, are irrational conservatives. But if useful design techniques have been developed, I think they should be used in space.

     As a matter of fact, you can get rougher with this. You can make a calculation on what would have happened in our air transport system last year if the philosophy of our ammunition people had been used in running it. If you do that, you'd find that we would have killed 4 million people last year. You will also find that the attrition of equipment and refurbishment costs are so high that it would cost you $10,000 for a ticket to anywhere. The only way out of the ammunition dilemma, surprisingly enough, is the kind of advance propulsion we're planning to talk about the next few days. You have to get enough margin into the propulsion so that you can have extra weight available both for use and reuse.

     An interesting interaction exists between the containment capabilities of gaseous fission systems and the cost of boost to orbit. Attempts are frequently made to show that high fuel consumption gaseous fission systems (either Orion or co-axial systems) would be acceptable after all, since the extra economic penalty which they incur compared to the cost of chemical boost to orbit is relatively small. This conclusion would obviously be strongly influenced by the wide spread of orbital costs mentioned.

• 

     Figure 14 shows typical interactions between containment of fuel and economics of boost to orbit. The point is obvious. Chemical take-off is not too bad. If chemical boost to orbit is tolerated at all, however, it must be of the economical "transportation" variety, or it will completely cripple the ability of gaseous fission engines to explore economically the solar system. This to me is the real challenge of advanced propulsion. This is also why I think there's a tremendous interaction between the engine and vehicle. It is not just a matter of sitting down with a specific impulse, and making one simple performance calculation. The big gain is made by the interaction of the engine and the development of transportation techniques.

• 

     Figure 15 is a sketch of one result of using a gaseous fission engine to power a reusable spaceship. We all, by now, expect manned rockets to be hundreds of feet long. If drawn to the scale of Figure 15, Saturn V would be two pages long. It would have a little bit of payload on the front. If, however, we were to combine the kind of nuclear rocket engine we would like to have (running on water or ammonia rather than hydrogen) with a reusable structure for the entire ship, a possible result would be the ship shown. This is a typical case of about a million pounds gross weight with cargo weight on the order of 200,000 pounds.

     It turns out, not surprisingly, that for reasonable economy, large payload fractions, perhaps even more than 20 percent, are required. Note that 20 percent cargo at a density of 10 pounds per cubic foot (standard transport airplane practice) when combined with the required propellant results in a rocket vehicle with 60 percent of its length devoted to cargo and crew. The engine and propellant take up only a small portion at the rear, just like "Buck Rogers" has said it should all along.

     An interesting example of the change in design philosophy with such ships is in the matter of shielding. Indications are that about 20,000 pounds of shielding weight would be required. This is a severe penalty for most rockets. Since cargo itself is effective shielding material, however, by the simple expedient of never flying this ship with less than 10 percent cargo aboard, properly packaged, the shielding penalty is reduced effectively to zero.

     If such a ship were used without radiators on the engine but with hydrogen propellant, it would be able to generate about 80,000 feet per second. This is more than needed for a lunar round trip. We can hence examine the effect of this ship on a lunar run, performing like a normal transport airplane. Logical assumptions, as previously discussed, would lead to 50 flights a year, and the ship can carry 100 tons per flight. By maneuvering a little bit with that number, I concluded that one ship like this is equivalent to 300 Saturn V launches per year.

     This is the kind of thing that we're driving at. Incidentally, in our scientific operations in the Antarctic, we deliver to the Antarctic about 50,000 tons a year. Ten such ships, shuttling back and forth to the moon, could mount the same magnitude of operation on the moon as we mount in Antarctica. This is, to say the least, an interesting capability.

• Fission Products Always in Vapor Form
  • No Burn-Up Problem as in Rover or Power Supplies
• No Fission Products aboard upon Return
  • Never Overfly with Fission Product Load
  • Servicing Problem Not Great as in ANP Application
• Fission Product Load small in case of Accident
  • Megapound Thrust is Only Kilo (Not Mega) Ton Equivalent
  • Only Millirem Exposure if Accident above 5000 Ft.
• High Velocity Exhaust Jet
  • Velocity of at Least 100,000 ft/sec exceeds Earth Escape Speed
    • Most Products actually Ejected from Solar System
• Characteristics also affect Development Procedures
  • Time between Static Tests Reduced
    • Engine does not contain Products after Test
    • Fuel Element Fabrication Delays Avoided
  • May Not be a Logical Extension of Solid Cores
    • Progress is Rarely Logical
    • ICBM was not Extrapolation of Winged Aircraft

Figure 16

     Some points should be made about the safety of gaseous fission spaceships. Contrary to most opinion, a gaseous fission rocket is probably a lot safer to use than a solid core rocket. Several reasons for this are listed on Figure 16. The fission products are always in vapor form, so there is never a fuel element burnup problem if emergency atmospheric entry is necessary. On return trips, one need never overfly a city with a fission product load, since the products can be ejected into space and the landing made aerodynamically. After landing, of course, the ship is radioactive only to the extent that any material has been locally activated. This will be very small with proper material selection and is certainly far lower than when the fission product load is a permanent feature of the structure. Hence, the servicing problem would be nowhere near as great as with the case of aircraft nuclear propulsion.

     Furthermore, in case of an accident, the fission product load is always small. A million pounds of thrust is, after all, only 1/2 kiloton of thrust, and the actual fission products created are comparable to those from kiloton, not megaton, bombs. As a matter of fact, some interesting calculations indicate that an accident as low as 5,000 feet in the air yields almost no exposure on the ground due to the effectiveness of atmospheric dispersal of the small fission product load on board.

     In addition to the fact that the fission product load is small, one can do interesting things by recognizing the fact that the exhaust jet velocity is actually higher than solar system escape speed. If trajectories are properly programmed, once out of the Earth's atmosphere, most of the fission products ejected with the exhaust will be thrown completely out of the solar system. This is the one way of not contaminating space. Space, incidentally, is a doggone big place, and the contamination of small local radiation belts, atmospheres, or planets should not be confused with all of space. Even our Sun, which is continually making a real attempt at space contamination compared to any puny spaceship, has not succeeded to any great distance.

     Some of the characteristics which make a gaseous fission rocket different from solid core propulsion systems also result in development differences. There is a tendency for many people to believe that gaseous fission engines are a logical extension of Rover. At the risk of losing a number of friends, I would like to point out that they probably are not a logical extension of Rover. For instance, the time between reactor tests should be greatly reduced for gaseous engines. One need not fabricate fuel elements between tests, and does not have to live with fission products imbedded within the engine. The handling advantages in operation previously mentioned also extend to the engine development process and the development program should be a lot easier to run than that of a solid core engine. Gaseous fission engines may not be a logical extension of solid core engines at all.

     A good space engine can also have a profound effect on spaceship development cost. It is not just a question of a high degree of reuse, there is also the effect of making the ship abortable during any part of the flight.

     Transportation systems not only achieve very high reuse, they contain sufficient redundancy to permit flying with partial equipment failures, and also have the ability to abort successfully from any flight condition. It is this last capability which is very important to the development program of such ships.

     The savings in development cost of not losing ships continually is obvious, yet our ammunition thinkers are so used to the massive throwaway that they usually claim that recoverable equipment would be more expensive to develop since it is more complicated. This might be true of the recovery of marginal performing rockets, but would not be true of a properly designed reusable spaceship which would not be marginal with a gaseous fission engine.

     It simply is not possible to overemphasize this difference in development philosophy. Commercial transports are extensively tested, and much of the equipment refined by flight tests. They become reliable pieces of equipment for expenditures very small in space budget terms, because it is possible to test the equipment over and over for a reasonable expenditure. If a high performance propulsion system can permit us for the first time to pursue a space vehicle program with the very efficient development techniques of transport aircraft systems, we will be very remiss if too blind to even consider these techniques.

     The use of a reusable and abortable ship from the start of the development program can well have a profound interaction on engine development tests. This interaction will be enhanced greatly if the engine is a "tractable" engine. A "tractable" engine is one which has "benign failure modes." In other words, it does not explode catastrophically when it fails. There is an excellent chance that gaseous fission engines will tend to go out rather than explode when trouble occurs.

     If the engine is tractable, and the ship abortable, then flight failures consist mostly of unscheduled landings. The ship is then capable of testing the main engine without the extreme sensitivity to component malfunction which exists in ammunition development programs. This can be a very large leverage on total development costs. Clearly, there must be extensive ground testing of the engines. However, it may be much easier to arrange partial duration ground runs in enclosed areas than it is to arrange total duration ground runs. The total duration runs could then be performed in the ship.

     Many engine developments in the past have made extensive use of flying test beds when ground facilities were not adequate. The technique should not be ignored if the flight vehicles are able to reintroduce it. This is one of the examples of development techniques available with transportation devices which are not within the realm of experience of ammunition developers.

     The final point I wish to make is, I'm against logical progress.

     A common mistake in development thinking seems to be a tendency to relate the basic performance achieved by a device with its development difficulty. It seems so logical to assume orderly progress in development programs. Actually, many major programs are not a result of orderly progress. One of the most recent interesting examples is the development of the ICBM. These ballistic missiles penetrate to their targets at a Mach number of 25. Orderly progress would have dictated that we build first fleets of supersonic bombers, then fleets of hypersonic bombers. Only after that would we consider whether or not Mach number 25 penetrators were desirable.

     The fact is that Mach number 25 ballistic missiles are considerably easier to build than hypersonic bombers (and evidently Mach number 3 bombers). Their performance is attained in a different manner, with different engines (not breathing air) and in a different flight region (out of the atmosphere). They are not a Mach number 25 airplane.

     The gaseous fission spaceship has many analogous elements. It is easy to achieve 500,000 fps out in space, as long as the engine is capable of it. A ship which never carries fission products aboard need never fight the safety problems of solid core nuclear rockets, or even the analogous problems of nuclear airplanes. An abortable transport rocket is a different development job than the building of larger ammunition. We must look at the gaseous fission ship in terms of its difficulty of development, not in awe of its possible accomplishments.

• 

• 

(ed note: This is an interesting design example from the always worth reading Bootstrapping Space blog by Chris Wolfe. It is mostly centered around estimating a mission delta V and sizing a propulsion system to fit, but his thought processes are interesting.)

All previous posts described all-chemical systems that could be built and operated profitably in the near term. This one focuses on electrical propulsion systems.
The defining features of most electric propulsion:
 - High efficiency (high Isp)
 - Low thrust
 - High power requirements
 - Long trip times
 - Long operating life

     I chose a specific paper (Frisbee, Mikellides) to examine since the authors thoughtfully included most of the interesting parameters for a reusable Nuclear Electric Propulsion (NEP) Mars cargo tug. I don't really dive into how to calculate this for yourself because the problem is quite difficult without modeling software.

     It all comes down to the details; the question of NEP vs. Solar Electric Propulsion (SEP) vs. Chemical depends on the specific mission goals and technologies used.

The summary:
23 tons dry mass for a nuclear-electric tug of ~6 MW thermal / 1.2 MW electric
64 tons cargo capacity from low Earth orbit to Phobos-Mars orbit
Just under 40 tons of water propellant for the outbound trip and another 7.2 tons acquired at Phobos for the return
2.2 years outbound, slightly less inbound
Two round trips between thruster refits, five round trips between reactor refits

---

Background:
(skip to the next section if you are already familiar with electric propulsion)

     The general idea is to use electrical power to dump energy into a propellant and then release it at very high speed.

     The simplest of these is Electrothermal.
     An electric current produces heat and the propellant is passed through it. First up is the resistojet, where a resistor somewhat like an incandescent lightbulb filament is heated and then the propellant is pumped past it. These are common devices in the RCS systems of satellites. Second is arcjet, which passes an electric arc directly through the propellant instead of through a resistor. These can reach higher Isp because they can heat the propellant beyond material limits for resistor elements.
     Efficiency is moderate (Isp of 500 to 1000, well above any realistic chemical system but easily an order of magnitude lower than the most efficient electrics). Preferred fuels are low atomic weight with no particulates (hydrogen, water, ammonia, hydrazine). Thermal power is most efficient when there are few degrees of freedom for the molecules so more of the energy can be applied to macroscopic motion, so hydrogen or hydrogen-rich propellants are ideal.

     Next up is Electrostatic.
     These are the 'traditional' ion thrusters, NSTAR, Hall effect, etc. The propellant is ionized by a strong electrostatic field (with some variations) and then the ions are accelerated by a negatively-charged grid or electrode. An electron gun is used to neutralize the electric charge of the ion beam and keep the spacecraft electrically neutral.
     Efficiency is high. with Isp ranging from 1000 to 5000 for most designs and a few reaching as high as 10,000. Preferred fuels are high atomic weight gases or elements with a very low first ionization energy; argon and xenon are frequently used. Sometimes iodine or metals like tin, magnesium or sodium are used, while caesium and rubidium can be used in a FEEP thruster.

     Lastly, Electromagnetic.
     These are the 'new generation' plasma thrusters like VASIMR, MPD, PIT, etc. Note the idea is not new, it's just that these technologies have been getting a lot of press lately. In fact, a pulsed plasma thruster (from this family) was the first electric thruster flown in space. In these devices the propellant is ionized by arc discharge, microwave heating or other means and accelerated by a magnetic field rather than an electrostatic grid.
 Efficiency is typically variable but can be very high, with an Isp of 1000 to 30,000 (most commonly about 1500 to 6000). Preferred fuels are the same as electrostatic thrusters for the most part, with lithium making an appearance in MPD thrusters.


     All three families have been in use for decades, while each family has relatively recent members pushing the limits. All share common fundamental physics with regard to their efficiencies. Sources of loss are in the power processing units, ionization energy of the propellant, dissociation energy of the propellant if it is a molecule rather than a pure element, ion impacts with the grid or body and in the charge density and geometry of the exhaust.

---

Application:

     Of key concern for a reusable vehicle is the propellant should be available in space. Xenon, argon and nitrogen are available in the atmosphere of Mars. Small amounts of nitrogen may be available in lunar cold traps or bound in soils of Ceres and many C-type asteroids. Water appears to be widely available. Alkali metals would be available on the Moon and in most asteroids.

     Another critical factor is that the thruster system should be reliable over the long term.  Electrostatic systems have already demonstrated very long operational lifespans in the range of 5 to 10 years of active thrust. Electromagnetic systems don't yet measure up in demonstrated lifespan, but that is mainly because most systems of this type are low Isp / high thrust RCS components designed for relatively short operating times rather than main drive units designed to last a decade. As a practical engineering concern, there are difficult challenges in improving lifespan of systems with exposed electrodes or grids. Designs without grids, whether electrostatic or electromagnetic, have the potential for decades-long operation.

     A practical system will be able to service a cargo route in a reasonable time. To illustrate this, let's dive into a paper on a pulsed inductive thruster proposal that includes payload transits to Mars, Saturn and solar escape. A reusable cargo tug with about 64 tons of payload and a 2.2-year Mars transit would require about 2 megawatts of electricity and about 275 tons fueled mass at departure. Propellant would be plain water, though ammonia works as well. If propellant supplies are available at Mars then we require only 1.2 megawatts and 165 tons fueled mass. Of that, the tug itself is just under 23 tons. Enough spare components would be included to make two round trips between thruster refits. A refit would mean swapping out the entire thruster pod with a fresh assembly, quick and easy.
     2.2 years is a long trip, but it is right at the synodic period for Earth and Mars. A cargo tug launched at one opportunity would arrive just in time to go / no go the crew launch at the next opportunity. Part of the problem is that a low-thrust vehicle has to produce about 16 km/s of dV for this trip, far more than a high-thrust chemical system requires thanks to the Oberth effect. Assuming an Isp of 6000 seconds the fuel mass ratio is 23.81% or 39.285 tons of water. An empty return trip would require 7.2 tons of propellant. Put another way, each ton of fuel delivers 1.6 tons of payload. Contrast that with my chemical tug's ratio of 0.8 tons of payload per ton of fuel and you can see the advantage; twice the mass delivered for the same quantity of fuel. Of course, the two approaches trade off costs between dry structure and fuel; electric propulsion is not automatically better but it certainly lets you do a lot more with the same starting mass.

     Power can come from two sources, solar or nuclear. The tug described above is nuclear; its reactor should be good for five or possibly six round-trips out of the box but could be designed for a much longer lifespan of 40 to 50 years with a bit more mass. This is partly because the same propulsion system is intended for exploration missions to the outer planets, where solar power is minimal.

     A Mars cargo tug could certainly use solar power instead, with a lifespan in the 20 to 30 year range and less complicated refitting / disposal. The ~19 tons of reactor, radiators and conversion hardware would be replaced by very large solar panel arrays of about 30 tons, 3.0 MW at Earth beginning-of-life yielding 1.2 MW at Mars end-of-life, with a whole-system specific power of 100 W/kg, 20-year useful life, 20% degradation and 50% of Earth-normal power available at Mars. If Photovoltaics (PV) refits are available every two or three trips then the allowance for degradation can be reduced to perhaps 10%, saving about 3.5 tons.

     Reactors also degrade over time as their fuel decays; some designs such as traveling wave, drum reflector and pebble bed can level the power output over time by only burning part of the nuclear fuel load at any one time. These designs can be life-extended by including more fuel during construction and / or by replacing fuel elements. Since the nuclear fuel is only a tiny fraction of a reactor's mass, this life extension adds very little mass to the overall system. This is the same reason why increasing a reactor's power takes less mass than increasing a solar panel array's power; for low power outputs the solar panels are nearly always lighter due to the reactor's heavy power conversion equipment and shielding, but as the power output grows the reactor eventually beats PV. As you can see from the above example, 1.2 megawatts at Mars after 20 years is firmly in nuclear territory given current state of the art solar performance. Even so, I would bet there is room in the design space for solar PV to be competitive at this power level, design life and solar distance.

---

Future work:

     I think the next step is to work up some EML2 to Phobos tether-capture cargo runs and see how they compare to the baseline LEO to LMO mission. Keeping used nuclear reactors out of LEO is a good idea, as is keeping large solar arrays out of the Van Allen belts. I'll try for an electric interplanetary tug with similar payload to my chemical tug. These would have the added bonus of providing abundant power while parked; there may be a case for a set of tugs such that one is always at Phobos providing megawatt-scale electrical power.

• 

• 

     I need to continue on the topic of electric propulsion. The previous post was a lot of words but not a lot of meat. I felt it was too weak to stand alone, particularly as a part of this series where I am trying to focus on a realistic near-term plan for cargo transport. If you are interested in more background information I'd start with the Wikipedia page on electric propulsion and follow up with a look at the Atomic Rockets engine page. Another good look in the context of interplanetary travel is this paper (Hellin), while a deep look at relevant equations can be had in this paper (Keaton).

     One interesting result is a general rule to find required thrust given average acceleration. Google failed me on finding an exact solution, but it looks like there is a simple approach that is within 1% of the target value.

     I eventually settled on a design massing 33.4 tons, 1.6 MW solar-electric, Isp 6,000 and 40 N thrust using PIT thrusters with water propellant.

     Let's look at an electric tug with payload comparable to my reference tug, both a solar PV and a nuclear version. The main routes for this vehicle will be between LEO, GEO, EML1/2 and Mars orbit. Unlike the chemical tug we can't get much out of the Oberth effect, so the delta-V requirements are higher. Just like the chemical tug, the LEO to EML1 leg has the highest dV requirements (about 7km/s), so if we design for that case then the other trips will be faster, carry more cargo or burn less propellant.

     A key design factor here is trip time. If we throw enough power at the problem we can get to EML1 in the same amount of time as a chemical rocket, but that is a poor use of the mass. We need to decide how long we are willing to wait for the cargo and design enough thrust into the ship to make the trip in that span. I'm going to suggest four weeks to EML1 as a reasonable compromise, so let's see the consequences of that choice.

---

Estimating thrust requirements for average acceleration

     To apply 7km/s of delta-V in 28 days we need to make an average acceleration of 2.89 mm/s. To allow some wiggle room let's assume we can only thrust 90% of the time, meaning now we need 3.22 mm/s. Since this is our average acceleration, we need to find either the initial or final acceleration to find the thrust of the propulsion system. To do that we will first need to know the vehicle's propellant mass fraction, so let's take a few test cases at Isp of 3000, 6000 and 10,000.

     Propellant mass fraction (Mf) is equal to 1 - e ^ (- dV / Ve), where dV in this case is 7000 m/s and Ve is Isp * g. See the rocket equation page for more details.
Isp 3000 -> Mf of 0.21175
Isp 6000 -> Mf of 0.11217
Isp 10000 -> Mf of 0.06890

     It would be nice if the average acceleration also matches up with the midpoint of fuel consumption, but somehow I doubt it. Let's find out.

     Given a dry mass of, say, 10 tons and an Isp of 3000, the fueled mass is ( 1 / ( 1 - Mf ) ) * dry mass, or 12.686 tons. When half of the fuel is burned the craft masses 11.343 tons. The target acceleration is 0.00322 m/s, so the required thrust is 36.52 newtons. Thrust is mass-flow (mdot) times exhaust velocity, so mdot is 1.2414 grams per second. That rate of propellant consumption would require 25 days to empty the tank, or 27.825 days after accounting for our 90% duty cycle.

     That surprises me. It's not exact but it is close enough for exploratory work. The case of 10,000 Isp works out to 27.947 days, so it looks like this general rule is valid across a fair range of Isp values. I also spot-checked some different mission dV values and found similar agreement, always within 1%. If anyone out there knows of an exact solution I would love to hear it.

     To calculate this yourself you need your mission dV, Isp, thrust duration and a test mass. If you set the test mass to 1kg (or 1t) then you can find a multiplier to use for different dry masses. The relationships are linear.

The required acceleration a will be dV in meters per second divided by thrust duration in seconds.
First, find fuel mass fraction, which is 1 - e ^ (- dV / Ve).
Convert to dry mass fraction Md, which is -Mf + 1
Convert to the 'gear ratio', which is simply 1 / Md
Multiply by dry mass M1 to get fueled mass M0 and note this value.
Find the fuel mass by taking M0 - M1 and note this value.
Find the 'halfway point', which is half the fuel mass plus M1; let's call this Mh.
Find the thrust F, which is Mh * a. This is the value you are looking for.
Find mdot, which is thrust divided by exhaust velocity, or F / (g * Isp ).
Find the real thrust duration, which is fuel mass divided by mdot. This should be within 1% of your stated thrust duration; if it is then the average acceleration value is accurate enough to use.

     If you have a known spacecraft (known dry mass and fuel mass, known thrust), you can use thrust divided by (dry mass plus half the fuel mass).

---

Electric tug design

     To align with the chemical tug, let's target a payload of 40 tons from LEO to EML1. Note that EML2 is a better target, but for purposes of comparison I'm using the LEO to EML1 trip as the most costly trip in the set. As mentioned above, we need to deliver in 28 days or provide an average acceleration of 3.22 mm/s. I don't have an exact solution, so I can't solve the problem in a single step. That's fine; spacecraft design is an iterative process.

     Let's assume an electric thruster at Isp = 6000 and mission dV of 7000 m/s. Also assume a one-way trip (meaning fuel is available at both endpoints). Power alpha is assumed to be 18 kg/kW, whether that be nuclear or long-life solar. Thrusters will be the NuPIT design shown in the last post, using the design values for the 5 N, 200 kW unit at 2.75 kg/kW (550 kg per thruster).
     As a first guess let's try eight thrusters, 1.6 MW. That's 33.2 tons, for a dry mass without tanks of 73.2 tons. We will need approximately 10 tons of liquid water propellant; using a tankage fraction of 2% would be reasonable in this case, so tack on 200kg for tanks for a total dry mass of 73.4 tons. Actual propellant load is 9,273 kg, so tankage is sufficient. The half-fueled mass is 78,037 kg and approximate average acceleration is 0.513 mm/s. We're not even close. Trip time would be 157.9 days, or 2.3 one-way trips per year.
     Maybe 20 thusters / 4 MW? Power alpha would improve to about 15, yielding 71 tons of power and propulsion. 40 tons of payload and perhaps 0.4 tons of tankage gives a dry mass of 111.4 tons, fuel mass of 14.1 t and average acceleration of 0.844 mm/s. This is clearly not going our way. Trip time would be 106.3 days, or 3.4 one-way trips per year.
     Let's aim much higher, 50 thrusters / 10 MW. Power alpha would continue to improve to about 14, yielding 167.5 t of power and propulsion. 40 tons of payload and 0.6 tons of tankage gives a dry mass of 208.1 tons, 26.29 tons of fuel and an average acceleration of 1.13 mm/s. Trip time would be 71.7 days or 5.1 one-way trips per year.

     Clearly, short trip times require increasingly absurd power levels. Matching the payload size of a chemical thruster with the 1.6 MW version means only making one round-trip per year. In fact, looking at that version of the ship, if we eliminate the payload entirely the highest acceleration the ship can make is 1.2 mm/s on its last gasp of propellant. Since fuel mass, dry mass, power and thrust are all linear relationships* that means no matter how we scale up the ship it can never get better than this. (The power system alpha does actually get better as we scale up, but moving a ship that masses several times your payload is inefficient and extremely expensive.)

     One thing we can do is increase the thrust of each propulsion unit, which usually means decreasing the Isp significantly. Let's look at a VASIMR thruster for comparison, since I have some data on performance at different Isp levels handy. A VASIMR thruster at 200 kW and 6000 Isp produces about 4.75 N of thrust, a fairly close match to the NuPIT. We need about six times that thrust (28.5 N), which occurs right at an Isp of 1000. That would bring the 8-thruster 1.6 MW vessel up to about 230 N of thrust. However, dropping the Isp so dramatically brings the fuel fraction just over 50%. That pushes our dry mass up to 75t, fuel mass to 78.1t and nets us only 2.0 mm/s average acceleration. It's a 40.5 day trip or 9 trips per year, but now we are burning more fuel than the chemical tug thanks to our drastically higher dry mass. Still no net benefit to be had.

---

Putting the tug to work

     Let's look at what an electric tug actually saves: propellant. In a fully functional ecosystem of cis-lunar services propellant is fairly plentiful. The speed, convenience and throughput of chemical vehicles far outweighs the efficiency of ion vehicles in this environment. Where an electric tug shines is in the buildup phase, where all of the propellant is coming from Earth. The tug would save money during a critical part of the project. What that means is we do not need to survive dozens of Van Allen belt transits over two decades, we just need to make a reasonable number of trips over two or three years. We also don't need to standardize on the same payload sizes as the chemical tug, nor do we need to make trips in 1 month. I would say that using the same power system alpha for the solar version as I do for the nuclear version is very pessimistic; these vessels would not need to function at Mars orbit, though they do need significantly thicker front-glass shielding on the panels than other craft.

     So, a lunar ISRU plan would still start with a single chemical tug / lander as described in part 1. Using performance for the detailed reference tug, a 15-ton package can be delivered from LEO direct to the lunar surface. This will be 12.4 tons of ISRU equipment and 2.6 tons of spares (2.1 year supply). Refilling tug 1 will take 6 months, after which it can deliver 33 tons to EML1.
     In the meantime, an electric tug (call it tug A) will deliver a 9-ton fuel depot (135 ton capacity) to EML1. Let's use our 40 N / 1.6 MW / 33.4t / 6000 Isp vehicle from above. It does the job in about 92 days, which means there is a window of three months after the launch of the first ISRU package to get the 33.4t tug, 5.36t propellant and 9t payload into LEO.
     At the first lunar launch, 33 tons are delivered to the EML1 depot. Tug A will collect this and head to LEO, taking 127 days and consuming 7.45t of water. During this trip a LEO depot is launched, identical to the one at EML1. The tug turns back around and heads for EML1, taking 4.22t of water for the return trip and leaving 21.33t of cargo in LEO. This could be a mix of surface samples and water as desired. Let's assume five tons are samples and the rest is fuel.
     The return trip takes 72 days, during which tug 1 will have delivered another 33 tons to the EML1 depot. Tug A repeats its performance, returning to LEO with a full load of 23.78t water and another 5t of samples. At this point we are at 598 days elapsed since start of ISRU operations, which should be enough time to settle on and construct additional hardware to expand the lunar surface capacity.
     This is significantly longer than the all-chemical scenario and has an IMLEO of 129.71 tons, within a few tons of all-chemical. Hardware costs are higher since more of the mass is spacecraft and much less of it is fuel. The main benefit is that schedule pressures are greatly reduced; final design, construction and testing of the second round of ISRU plant is allowed more than a year and a half of time rather than two months. More operational data is available and the tolerance for mistakes or inefficiencies is higher. Another benefit is that this profile includes depots in LEO and at EML1; even if things do not progress beyond the first ISRU package the infrastructure is still useful for this and future projects.
     This baseline hardware could continue to deliver 21 tons of cargo to LEO every ~200 days for about a decade, eventually reaching 426 tons over 20 trips at a cost of 141 tons of Earth mass or a leverage of about 3 to 1. Things improve if we continue to expand, since about 44% of that mass was fuel to get the first ISRU plant in position; additional ISRU hardware is delivered using lunar propellant.

     The next phase would be to send more ISRU hardware. Tug A can pick up 33 tons at EML1, deliver 19.18t of net payload to LEO over 127 days, pick up a 17-ton package and head to EML1 in 109 days. All of the required propellant is lunar and picked up at EML1. Round trip time is 236 days (a bit under 8 months). The harvesting process run by tug 1 has a shorter turnover time of 6 months, so on average an extra 19 tons is accumulated at the depot. That's not quite enough to provide for a cargo landing, so tug A may not always be bringing a full load of cargo to LEO (meaning shorter round trips in practice).
     An alternative might be to use 12-ton packages that will fit into a Falcon 9 for cheaper launch costs; the delivery time for that is 98 days. If less cargo is returned to LEO then that trip time can be shortened as well; for example, 6 tons of return cargo plus round-trip fuel would make each leg of the trip take 98 days, or 196 days round-trip. Each 6.5-month trip would deliver another 10 tons of ISRU with two years of spares. Two electric tugs could deliver 80 tons of ISRU capacity in 26 months, roughly a single Mars synodic period. That would place 95 tons of ISRU with expected output of 950 tons of propellant annually at a cost of 9.5 tons of spares. Net propellant delivered to EML1 would be 505 tons annually, or could be 168 tons to LEO annually with chemical tugs. The annual demand for spares (both ISRU and depots) can be met in a single hardware run with minimal fuel costs, leaving 3/4 of the electric tug schedule open for assignments like delivering new chemical tugs or GEO debris retrieval (a mission that avoids the majority of the radiation belts and prolongs the tug's useful life).
     The total phase 2 IMLEO would be 134.4 tons, all of it hardware. Lunar mass to LEO during this period would only be another 48 tons since capacity is focused on buildup. This phase would run for 26 months, or a total of 46 months since first launch.
     Ongoing maintenance would require approximately 12 tons per year. The initial depots would be insufficient, so we need another 27t of hardware for fuel storage. If we rate all flight and depot hardware with a 10-year lifespan and pro-rate the replacement mass then we need an additional 12.2 tons annually (24.2t total). Depending on how the output is allocated, this could be considered an ongoing leverage of 23.5 tons in EML1 per ton IMLEO or 7.8 tons in LEO per ton IMLEO. Another way to look at it would be as a fuel supply for three manned Mars missions covering four synodic periods (104 months), or a full ISRU program length of 150 months (12.5 years). Overall Earth mass to LEO is then 511.15t to harvest 5,501.7 tons of gross lunar propellant, yielding 2,924.6 tons net lunar propellant at EML1. That's a gear ratio of about 5.7 to 1. If you are only interested in delivering fuel to LEO then you can net 972.9 tons, still a favorable 1.9 to 1 mass ratio. 511 tons is a lot of mass to launch, but only three payloads require a heavy lift vehicle: the initial chemical tug stack (62t fuel and 22t hardware, split across two Falcon H) and the two electric tugs (33.4t each, also requiring a Falcon H unless they can be built in parts and flown on two Vulcan launches). The remaining 360 tons would be delivered by 30 Falcon 9 launches, or by some combination of any price-competitive launchers with at least 12 tons of payload.

     Launch costs would be roughly $2.1 billion. Hardware would run another $6.7 billion (at $15m per ton). Operations might cost $125-$250 million. Call it a total of $9 billion over about 15 years (12.5 years of operation plus 2.5 years of r&d, manufacturing and testing). Overall cost of fuel at EML1 would be $3,094.44 per kg, about $3.1 million per ton and expected to decline to $0.9 million per ton in the long run. Savings are about the same as the all-chemical approach, a bit over $4.5 billion vs. NASA baseline. Additional savings could be realized by using the chemical tugs as cargo haulers to and from Mars as described in part 2, resulting in excess capacity that could be sold or used for other purposes. One of those purposes might be ISS reboost and water supply for life support. Another might be developing a significant water supply on the Moon for growing food, in support of manned missions.

• 

Ship mass and size

Full load mass and physical size depends upon assumptions about fuel mass ration, fuel bulk, etc.

Assumption:

Deadweight (inert mass)	1	17%
Cargo (payload)	2	33%
Fuel (propellant)	3	50%
TOTAL	6	100%

Note that total mass is three times the cargo capacity. As you can see, deadweight is the ship proper, structure, engines, anything that is not cargo or propellant.

With this assumption, the big freighters will have a fully loaded mass of 60,000 tons. The largest ships might be twice as big: 120,000 tons.

Our building cost is $500,000 per ton of cargo capacity, the mass assumption makes a building cost equal to $1 million per ton of deadweight. Annual service cost is $100,000 per ton of cargo capacity, the mass assumption makes the annual service cost equal to $200,000 per ton of deadweight. The starship hulls are not cheaper, but they can carry more cargo in proportion to their structural mass.

Type of ship	Cargo capacity	Purchase price
Large	20,000 tons	$20 billion
Medium	5000 tons	$2.5 billion
Small	1500 tons	$750 million

At $500,000 per ton of cargo capacity, largest giant freighter cost $20 billion to build, but it it has a cargo capacity of 200 Boeing 747 jets, and accounts for over one percent of whole fleet's cargo capacity all by itself. Small freighter costs $750 million, and has seven time the capacity of 747.

With a 30 year service life, the combined shipbuilding yards of the 12 planet trade network will turn out about 25 ships per year.

Hulls will last longer than 30 years but the equipment wears out and has to be replaced. Ships go back to the yards for an overhaul every decade or so, but eventually the cost of stripping everything and replacing it will exceed the value of the ship. Depending upon overhaul costs the shipyards may make more money on rebuilding than on constructing brand new ships. Some ships will stay in service for many decades. Others will be retained as the futuristic equivalent of naval hulks or the old passenger equipment that railroads use as work trains. Every big commercial space station will have a bunch of these old ships in the outskirts.

If modular design is taken to its limit, "ships" will have no permanent existence. Instead they will be assembled out of modules and pods specifically for each run, much like a railroad train. In that case, a ship's identity is attached to a service, not a physical structure. Example: the Santa Fe "Chief" was identified by a timetable and reputation, not a particular set of locomotive and cars.

---

• 

Starship Performance

The analysis up until now focused on money and economics. Businessmen only care about how long it takes to deliver the cargo and how much transport costs, they could care less about the scientific details of the ship engines. But authors care.

As with everything else, it all depends upon the assumptions. Your assumptions will be different, so feel free to fiddle with these and see what the results are.

Assumption: the time spent in FTL transit is zero (jump drive). For the FTL segment of the transit you can use whatever you want, as long as the details do not affect the analysis. The main thing is that the required time spent in FTL transit will add to the total trip time, and thus the number of cargoes a starship can transport per year.

Assumption: starships use reaction drives for normal space travel.

We know that the mass ratio is 2.0. So the Tsiolkovsky rocket equation tells us that the starship's total delta V will be the propulsion system's exhaust velocity times 0.69 (i.e., ln(2.0) ). Since starships accelerate to half their delta V, coast, then decelerate to a halt, their maximum speed is half their delta V, or exhaust velocity times 0.35 (i.e., ln(2.0) / 2). In practice you would accelerate up to a bit less than half their delta V in order to allow a fuel reserve in case of emergency.

It will be even less if the FTL drive happens to use the same type of fuel that the reaction drive does. Basically part of the fuel mass will have to be considered as cargo, not propellant, which will alter the ship's mass ratio.

Reaction drive	Exhaust velocity general rule
Nuclear powered Ion	~100 km/s
Fusion	a few thousand km/s
Beam core matter-antimatter	about 100,000 km/s ( 1/3 c )

We have assumed that the ship spends 27 days in route (with an instantaneous FTL jump), so the outbound and inbound legs are 13.5 days each (1.17 million seconds).

Assumption: the acceleration on each leg is constant. In reality at the same thrust setting the acceleration will increase as the ship's mass goes down due to propellant being expended. The thrust will probably be constantly throttled to maintain a constant acceleration. Makes it easier on the crew and easier on our analysis. The implication is that obviously the average speed will be half the maximum speed (which is half the delta V)

Reaction drive	Exhaust velocity general rule	Average speed	Outbound/ inbound leg distance	Acceleration/ deceleration
Advanced Ion or Early Fusion	400 km/s	130 km/s	75 million km (1/2 AU)	0.01 g
Advanced Fusion	10,000 km/s	5000 km/s	20 AU (Sol-Uranus)	0.44 g
Beam-core Matter-Antimatter	c	0.3 c	350 AU (x5 Pluto's orbit)	8 g !!!

These figures will be lower if time is consumed in FTL flight, maybe be only Terra-Luna distance

Propulsion system's thrust power is thrust times exhaust velocity, then divide by 2. To get the thrust, we know that thrust is ship mass times acceleration. The ship mass goes down as fuel is burnt. As a general rule for ship mass, figure that it only has 2/3rds of a propellant load. That is, multiply the total ship mass by 0.83. So our 120,000 metric ton ship would have a general rule mass of 120,000 * 0.83 = 100,000 metric tons (100,000,000 kilograms).

Reaction drive	Exhaust velocity general rule	Acceleration/ deceleration	Thrust	Thrust power
Advanced Ion or Early Fusion	400,000 m/s (400 km/s)	0.108 m/s (0.011 g)	1.08×107 N	2.16×1012 W (2 terawatts)
Advanced Fusion	10,000,000 m/s (10,000 km/s)	4.3 m/s (0.44 g)	4.3×108 N	2.15×1015 W (2,000 terawatts)
Beam-core Matter-Antimatter	3.0×108 m/s (c)	76.5 m/s (7.9g)	7.65×109 N	1.15×1018 W (1 million terawatts)

---

• 

Where does fuel come from and who does it get into the ship's fuel tanks? Easiest if it is obtained locally at the destination's solar system. The economics of interplanetary transport is same as interstellar (since we did a lot of work making interstellar a cheap as interplanetary).

if fuel from a gas giant at a distance comparable to Terra-Jupiter and round trip is to only take weeks, interplanetary tankers will need speeds of around 1000 km/s. So tankers will be almost as expensive as starships. If tankers use low speed (to make them cheaper), the round trip balloons to a year or more. To service the starship fleet's thirst for fuel, tankers will need to be huge or there will have to be a lot of them. Either way, fuel shipped from gas giants ain't gonna be cheap.

If we forgo interplanetary tankers and instead have starships make extra leg to the local gas giant to refuel, it will cost you more than you will save.

The alternative is shipping fuel up from destination planet. Yes, we know about how surface to orbit is "halfway to anywhere" in terms of delta V cost. But in order to colonize space at all, surface-to-orbit shipping cost will have to be cheap anyway. The industrialization of space will start with using space based resources, but eventually surface-to-orbit will have to be cheap or there is no rocketpunk future. Laser launch, Lofstrom loop, space elevator, something like that.

Assumption: surface-to-orbit shuttle economics are equivalent to current day airliner economics. Round trip to LEO and back is about two hours (not counting loading/unloading). With loading/unloading and maintenance, figure 4 flights a day. Implication is that a round trip passenger ticket is $250 and round trip freight service is $1000/ton (which is +10% added to interstellar transport costs)

Fuel is not round trip, it only goes from surface to orbit, but shuttles have to go orbit to surface in order to get the next load. You will have to streamline the process. High capacity pumps to minimize load/unload times, crew-less shuttle. You might be able to squeeze fuel lift cost to $500/ton. So if starships carry 1.5 tons of fuel per ton of cargo, surface-to-orbit fuel lift costs adds $750/ton to interstellar shipping cost.

So total surface-to-orbit overhead is $1000/ton + $750/ton = $1750/ton or 17.5%. This is an ouch but not a show-stopper.

---

Back to starships. How big are they?

Present-day maritime tonnage rule: 1 registered ton = ~3 cubic meters.

Assumption: 1 ton = 3 m³ applies to fuel and hull (e.g., crew quarters, engineering spaces, etc) as well as cargo. Therefore, if the absolutely hugest cargo starship in service has a cargo capacity of 40,000 tons (twice that of a large cargo starship), then:

Wet Mass

Payload mass to total mass ratio is 3. So wet mass is 3 * 40,000 = 120,000 tons

Starship Volume

1 ton of total ship mass = 3 m³ of volume. 120,000 * 3 = 360,000 cubic meters.

Volume of a sphere is 4/3πr³, so the radius of a sphere is ³√(v/(4/3π)) or

radius = CubeRoot( v / 4.189)

diameter = (CubeRoot( v / 4.189)) * 2

• 

Assumption: a "cigar-shape" for a spacecraft is a six times as long as it is wide, with the proportions indicated in the diagram above. The center body is a cylinder 1 unit in diameter (0.5 units radius) and two units high. The two end caps are cones of 0.5 units radius and 2 units high.

If the monstrous cargo starship is spherical, it would have a diameter of 88 meters. If it is cigar shaped then length = 300 meters and diameter of 50 meters.

A 1500 ton cargo capacity tramp freighter would have a wet mass of 4500 tons and a volume of 13,500 m³. Spherical shape would have a diameter of 30 meters, cigar shaped length = 100 meters long and diameter of 17 meters.

Modular ships dimension would be similar but a bit larger due to being assembled out of component parts.

When a spacecraft built for humans ventures into deep space, it requires an array of features to keep it and a crew inside safe. Both distance and duration demand that spacecraft must have systems that can reliably operate far from home, be capable of keeping astronauts alive in case of emergencies and still be light enough that a rocket can launch it.

Missions near the Moon will start when NASA’s Orion spacecraft leaves Earth atop the world’s most powerful rocket, NASA’s Space Launch System. After launch from the agency’s Kennedy Space Center in Florida, Orion will travel beyond the Moon to a distance more than 1,000 times farther than where the International Space Station flies in low-Earth orbit, and farther than any spacecraft built for humans has ever ventured. To accomplish this feat, Orion has built-in technologies that enable the crew and spacecraft to explore far into the solar system.

Systems to Live and Breathe

As humans travel farther from Earth for longer missions, the systems that keep them alive must be highly reliable while taking up minimal mass and volume. Orion will be equipped with advanced environmental control and life support systems designed for the demands of a deep space mission. A high-tech system already being tested aboard the space station will remove carbon dioxide (CO2) and humidity from inside Orion. Removal of CO2 and humidity is important to ensure air remains safe for the crew breathing. And water condensation on the vehicle hardware is controlled to prevent water intrusion into sensitive equipment or corrosion on the primary pressure structure.

The system also saves volume inside the spacecraft. Without such technology, Orion would have to carry many chemical canisters that would otherwise take up the space of 127 basketballs (or 32 cubic feet) inside the spacecraft—about 10 percent of crew livable area. Orion will also have a new compact toilet, smaller than the one on the space station. Long duration missions far from Earth drive engineers to design compact systems not only to maximize available space for crew comfort, but also to accommodate the volume needed to carry consumables like enough food and water for the entirety of a mission lasting days or weeks.

Highly reliable systems are critically important when distant crew will not have the benefit of frequent resupply shipments to bring spare parts from Earth, like those to the space station. Even small systems have to function reliably to support life in space, from a working toilet to an automated fire suppression system or exercise equipment that helps astronauts stay in shape to counteract the zero-gravity environment in space that can cause muscle and bone atrophy. Distance from home also demands that Orion have spacesuits capable of keeping astronaut alive for six days in the event of cabin depressurization to support a long trip home.

Proper Propulsion

The farther into space a vehicle ventures, the more capable its propulsion systems need to be to maintain its course on the journey with precision and ensure its crew can get home.

Orion has a highly capable service module that serves as the powerhouse for the spacecraft, providing propulsion capabilities that enable Orion to go around the Moon and back on its exploration missions. The service module has 33 engines of various sizes. The main engine will provide major in-space maneuvering capabilities throughout the mission, including inserting Orion into lunar orbit and also firing powerfully enough to get out of the Moon’s orbit to return to Earth. The other 32 engines are used to steer and control Orion on orbit.

In part due to its propulsion capabilities, including tanks that can hold nearly 2,000 gallons of propellant and a back up for the main engine in the event of a failure, Orion’s service module is equipped to handle the rigors of travel for missions that are both far and long, and has the ability to bring the crew home in a variety of emergency situations.

The Ability to Hold Off the Heat

Going to the Moon is no easy task, and it’s only half the journey. The farther a spacecraft travels in space, the more heat it will generate as it returns to Earth. Getting back safely requires technologies that can help a spacecraft endure speeds 30 times the speed of sound and heat twice as hot as molten lava or half as hot as the sun.

When Orion returns from the Moon, it will be traveling nearly 25,000 mph, a speed that could cover the distance from Los Angeles to New York City in six minutes. Its advanced heat shield, made with a material called AVCOAT, is designed to wear away as it heats up. Orion’s heat shield is the largest of its kind ever built and will help the spacecraft withstand temperatures around 5,000 degrees Fahrenheit during reentry though Earth’s atmosphere.

Before reentry, Orion also will endure a 700-degree temperature range from about minus 150 to 550 degrees Fahrenheit. Orion’s highly capable thermal protection system, paired with thermal controls, will protect Orion during periods of direct sunlight and pitch black darkness while its crews will comfortably enjoy a safe and stable interior temperature of about 77 degrees Fahrenheit.

Radiation Protection

As a spacecraft travels on missions beyond the protection of Earth’s magnetic field, it will be exposed to a harsher radiation environment than in low-Earth orbit with greater amounts of radiation from charged particles and solar storms that can cause disruptions to critical computers, avionics and other equipment. Humans exposed to large amounts of radiation can experience both acute and chronic health problems ranging from near-term radiation sickness to the potential of developing cancer in the long-term.

Orion was designed from the start with built in system-level features to ensure reliability of essential elements of the spacecraft during potential radiation events. For example, Orion is equipped with four identical computers that each are self-checking, plus an entirely different backup computer, to ensure Orion can still send commands in the event of a disruption. Engineers have tested parts and systems to a high standard to ensure that all critical systems remain operable even under extreme circumstances.

Orion also has a makeshift storm shelter below the main deck of the crew module. In the event of a solar radiation event, NASA has developed plans for crew on board to create a temporary shelter inside using materials on board. A variety of radiation sensors will also be on the spacecraft to help scientists better understand the radiation environment far away from Earth. One investigation called AstroRad, will fly on Exploration Mission-1 and test an experimental vest that has the potential to help shield vital organs and decrease exposure from solar particle events.

Constant Communication and Navigation

Spacecraft venturing far from home go beyond the Global Positioning System (GPS) in space and above communication satellites in Earth orbit. To talk with mission control in Houston, Orion’s Orion will use all three of NASA’s space communications networks. As it rises from the launch pad and into cislunar space, Orion will switch from the Near Earth Network to the Space Network, made possible by the Tracking and Data Relay Satellites, and finally to the Deep Space Network that provides communications for some of NASA’s most distant spacecraft.

Orion is also equipped with backup communication and navigation systems to help the spacecraft stay in contact with the ground and orient itself if it’s primary systems fail. The backup navigation system, a relatively new technology called optical navigation, uses a camera to take pictures of the Earth, Moon and stars and autonomously triangulate Orion’s position from the photos. Its backup emergency communications system doesn’t use the primary system or antennae for high-rate data transfer.

• 

A geodetic airframe is a type of construction for the airframes of aircraft developed by British aeronautical engineer Barnes Wallis in the 1930s (who sometimes spelled it "geodesic"). Earlier, it was used by Prof. Schütte for the Schütte Lanz Airship LS 1 in 1909. It makes use of a space frame formed from a spirally crossing basket-weave of load-bearing members. The principle is that two geodesic arcs can be drawn to intersect on a curving surface (the fuselage) in a manner that the torsional load on each cancels out that on the other.

Aeroplanes

• 

The earliest-known use of a geodesic airframe design for any aircraft was for the pre-World War I Schütte-Lanz SL1 rigid airship's envelope structure of 1911, with the airship capable of up to a 38.3 km/h (23.8 mph) top airspeed.

The Latécoère 6 was a French four-engined biplane bomber of the early 1920s. It was of advanced all-metal construction and probably the first aircraft to use geodetic construction. Only one was built.

Barnes Wallis, inspired by his earlier experience with light alloy structures and the use of geodesically-arranged wiring to distribute the lifting loads of the gasbags in the design of the R100 airship, evolved the geodetic construction method (although it is commonly stated, there was no geodetic structure in R100). Wallis used the term "geodetic" to apply to the airframe and distinguish it from "geodesic" which is the proper term for a line on a curved surface, arising from geodesy.

The system was later used by Wallis's employer, Vickers-Armstrongs in a series of bomber aircraft, the Wellesley, Wellington, Warwick and Windsor. In these aircraft, the fuselage was built up from a number of duralumin alloy channel-beams that were formed into a large framework. Wooden battens were screwed onto the metal, to which the doped linen skin of the aircraft was fixed.

The metal lattice-work gave a light structure with tremendous strength; any one of the stringers could support some of the load from the opposite side of the aircraft. Blowing out the structure from one side would still leave the load-bearing structure as a whole intact. As a result, Wellingtons with huge areas of framework missing continued to return home when other types would not have survived; the dramatic effect enhanced by the doped fabric skin burning off, leaving the naked frames exposed (see photo). The benefits of the geodesic construction were partly offset by the difficulty of modifying the physical structure of the aircraft to allow for a change in length, profile, wingspan etc.

Geodetic wing and fin structures—taken from the Wellington—were used on the post-war Vickers VC.1 Viking, though with a new fuselage and metal-skinned.

• 

•