Abstract

     The Revolutionary Aerospace Systems Concepts (RASC) team, led by the NASA Langley Research Center, is tasked with exploring revolutionary new approaches to enabling NASA to achieve its strategic goals and objectives in future missions. This paper provides the details from the 2004-2005 RASC study of a point-design that uses a high-power nuclear electric propulsion (NEP) based space transportation architecture to support a manned mission to Mars. The study assumes a high-temperature liquid-metal cooled fission reactor with a Brayton power conversion system to generate the electrical power required by magnetoplasmadynamic (MPD) thrusters. The architecture includes a cargo vehicle with an NEP system providing 5 MW of electrical power and a crewed vehicle with an NEP system with two reactors providing a combined total of 10 MW of electrical power. Both vehicles use a low-thrust, high-efficiency (5000 sec specific impulse) MPD system to conduct a spiral-out of the Earth gravity well, a low-thrust heliocentric trajectory, and a spiral-in at Mars with arrival late in 2033. The cargo vehicle carries two moon landers to Mars and arrives shortly before the crewed vehicle. The crewed vehicle and cargo vehicle rendezvous in Mars orbit and, over the course of the 60-day stay, the crew conducts nine-day excursions to Phobos and Deimos with the landers. The crewed vehicle then spirals out of Martian orbit and returns via a low-thrust trajectory to conduct an Earth flyby. The crew separates from the vehicle prior to Earth flyby and aerobrakes for a direct-entry landing.

Introduction

     This paper details a Revolutionary Aerospace Systems Concepts (RASC) study investigating a highpower nuclear electric propulsion (NEP) space transportation architecture to support a manned mission to Mars. The RASC project, led by the NASA Langley Research Center, is tasked with exploring revolutionary new approaches to enabling NASA to achieve its strategic goals and objectives in future missions. For this study, two vehicle concepts were designed, both using a high-power NEP system with Brayton power conversion and magnetoplasmadynamic (MPD) thrusters. The first vehicle is the Mars Transfer Vehicle (MTV) which carries the crew from the Earth to Mars and back again. The second vehicle, the Cargo Transfer Vehicle (CTV), delivers additional cargo necessary for the mission from the Earth to Mars.

     This paper details one of the four space transportation architectures selected by the 2004-2005 RASC Mars Obiter Study for analysis. The other three investigated were nuclear thermal propulsion (Borowski, Packard, and McCurdy, 2006), NEP with Rankine power conversion, and chemical propulsion. In order for the architectures to be compared across an even playing field, all four started with the same mission and payload assumptions. The mission consisted of a split profile with the cargo elements sent out on one vehicle and the crew sent out on a second vehicle. Each transportation architecture in the RASC study assumed the same cargo and crew payloads. These study requirements led to a mission that was not optimized specifically for an NEP system.

Vehicle Configurations

Mars Transfer Vehicle (MTV)

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     In order to provide the required artificial gravity for the crew during the Trans-Mars Injection (TMI) outbound and Trans-Earth Injection (TEI) inbound trajectory legs, the Mars Transfer Vehicle was configured to allow a rotation about the center of gravity. The crew is located in an inflated Transportation Habitat (TransHab) at one end of the NEP vehicle while the Brayton power conversion system and the nuclear reactors are located at the other end. To minimize the translation of the center of gravity over the mission, the LH₂ tanks are located at the center of the vehicle configuration. The MTV uses two reactors, each providing 5 MW_e, and a total of four Brayton power conversion units. There are two thruster arms with four 2.5 MW_e MPD thrusters (two operational, two spare) on each arm. Each thruster arm has a radiator to reject heat from the power processing units (PPU). The total planform area of the PPU radiators is 136.7 m² (273.4 m² effective radiating area). Six LH₂ tanks that are 7.6 m in diameter and 19 m long occupy the middle truss section of the vehicle and store the 279.4 MT of propellant. The main radiator is comprised of two sections of double-sided flat panels attached to the center truss structure on either side of the propellant tanks due to center of gravity requirements. The total planform area of the main radiator is 2722 m² (5444 m² effective radiating area). The MTV is 182 m long and must be assembled in orbit. The configuration of the MTV is shown in figure 1.

Cargo Transfer Vehicle (CTV)

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     The Cargo Transfer Vehicle is modeled after the NEP configuration used in the 2002 RASC Callisto mission entitled HOPE (McGuire, et al., 2003, and Borowski et al., 2003). Since the cargo vehicle does not require the artificial gravity spin, the propellant tanks are located at the far end from the reactor to prevent splitting up the radiator into two sections. This avoids having the hot heat-rejection fluid routed around the cryogenic tanks, as is required in the MTV configuration. Like the MTV, the CTV has doublesided radiator panels attached to the central truss structure of the vehicle. The total planform area of the main radiator is 1361 m², which provides 2722 m² effective radiating area. The CTV uses one reactor and two Brayton power conversion units to provide 5 MW_e power. The four 2.5 MW_e MPD thrusters are mounted on the outside of the truss section that contains the propellant tanks with two thrusters, one of which is a spare, on each side. The total planform area of the PPU radiators is 273.4 m² (546.8 m² effective radiating area). The CTV only has two of the 7.6 m diameter, 19 m long LH₂ tanks, storing the 63.9 MT of propellant. The CTV is 127 m long and, like the MTV, must be assembled in orbit. The configuration of the CTV is shown in figure 2.

Assumptions

Mission Assumptions and Outline

     The RASC Mars Orbiter mission was configured as an opposition class (short stay) Earth-to-Mars round-trip mission. A crew of six is deployed to Mars, but does not perform any Mars surface operations. Rather, they perform two nine-day excursions to Phobos and Demos before returning home to Earth. The total stay-time in Mars orbit is 60 days. The components of the NEP stages of both vehicles are launched on heavy-lift Magnum expendable launch vehicles (ELVs) and assembled in a circular Low Earth Orbit (LEO) at 1000 km altitude and 28.5° inclination. The Magnum is assumed to be capable of delivering 80 MT into LEO in a payload shroud 7.5 m wide by 30 m long.

     The CTV conducts a spiral escape from Earth and follows a low-thrust trajectory to Mars to predeploy two moon landers (for landing on Phobos and Deimos) in Mars orbit prior to the crew’s arrival. After assembly and checkout, a second NEP stage with the TransHab begins the spiral escape from LEO. After the NEP stage has cleared the Van Allen belts and is ready to escape Earth, the crew is launched on a smaller ELV (Delta IV Heavy class) in an Earth crew return vehicle (ECRV) and docks with the NEP stage in a high orbit. At this point, the mated NEP stage with the inflated TransHab and ECRV is referred to as the MTV. The MTV uses the reaction control system (RCS) thrusters to spin the MTV end-over-end upon Earth escape to provide artificial gravity (38 percent of Earth gravity, equal to Mars gravity) to the crew in the TransHab module. The MPD thrusters provide for “side thrusting” by thrusting along the axis of rotation. Once the MTV has reached Mars space, the vehicle performs a spin down maneuver, and the MTV spiral captures into the same Mars orbit as the CTV.

     After 60 days of Mars orbit operations, the MTV spiral escapes from Mars orbit and follows a lowthrust trajectory back to Earth. During the heliocentric portion of the flight, the RCS thrusters induce another end-over-end spin for artificial gravity (38 percent of Earth gravity) for the crew in the TransHab. At Earth arrival, the ECRV separates from the MTV with the crew onboard to perform a direct-entry aerobrake and parachute landing on Earth.

Payload Assumptions

     With this mission architecture, the cargo (moon landers) is sent out on a separate vehicle than the crew. The crew only carries enough supplies and cargo to last them through the TMI leg, the 60-day stay, and the TEI leg of the trip. All cargo necessary to carry out moon-landing operations at the destination is sent out on the CTV. The CTV payload consists of two moon landers designed by a team led by the NASA Langley Research Center as part of this RASC study. These landers are designed to take three crew on nine-day excursions to the surfaces of Deimos and Phobos, and then return to Mars orbit to rendezvous with the MTV.

     The MTV payload consists of an inflatable TransHab and an ECRV. The TransHab is similar to the TransHab design from the Human Exploration and Development of Space (HEDS) design reference mission 4.0 study (Joosten, 2002). The TransHab mass includes enough consumables for a 545-day round-trip mission. Any missions with total trip times longer than 545 days must add an additional 2.45 kg/person/day to the dry-mass allocation. The crew are onboard the MTV for a total of 612 days, so this adds 984.9 kg of consumables to the TransHab for this study. The mass of the TransHab also includes approximately 1900 kg of water for radiation protection and 400 kg for the environmental control and life-support system. The ECRV carries the crew during the final aerobrake for an Earth landing at the end of the mission. Table 1 shows the masses for each of the piloted and cargo payload elements as set by the RASC study. These masses already contain the appropriate contingency for each item, so no additional contingency was added to the payloads in this study.

TABLE 1.—RASC 2004 PAYLOADS

Element	Mass (MT)	Vehicle
TransHab: includes food for 545 days, 6 crew	35.0	MTV
Earth crew return vehicle (ECRV)	7.0	MTV
ECRV docking structure	8.0	MTV
Two moon landers for 9-day missions	42.5	CTV

Power System

     This study assumes a high-temperature, liquid-metal, fission reactor with a Brayton power conversion system to generate the electrical power required to supply the MPD thrusters. The reactor was based on an advanced version of the early reactor concept for the Jupiter Icy Moons Orbiter study. The fission reactors use liquid-metal coolant loops, which operate at a temperature of about 1600 K, in order to represent “mid-term” technology (Mason, 2001), consistent with the 2033 mission timeframe. Each reactor coolant loop transfers heat to the Brayton system’s working fluid via a heat source heat exchanger (HSHX), producing a Brayton turbine inlet temperature of 1500 K. The Brayton unit includes a recuperator to improve system efficiency by pre-heating the working fluid from the compressor outlet with the turbine exhaust before it reaches the gas cooler. The recuperator reduces the heat load of both the gas cooler and the HSHX, which in turn reduces the size of the radiators and the reactor. The heat rejection system uses a pumped NaK working fluid to remove heat from the Brayton working fluid via the gas cooler and transfers that heat to the radiator panels via water heat pipes. A turbine inlet temperature of 1500 K requires a very high-temperature turbine blade material (possibly ceramic) or active cooling of the blades, and a reactor temperature of 1600 K necessitates the use of refractory metals or other high temperature material for the reactor. A schematic of the Brayton power conversion system is shown in figure 3.

     The MTV uses two reactors sized to provide 5 MW_e net electrical power, each. Neutron interactions between the two reactors were not considered. The cargo vehicle only requires one reactor sized to provide 5 MW_e net electrical power. The component masses and radiator areas for both vehicles are presented in table 2. The reactor system includes the radiation shield, which is composed of layers of tungsten (gamma shield) and lithium hydride (neutron shield). The MTV’s shields are much heavier than the CTV’s due to the crew’s more stringent radiation limits. The radiator is double-sided, so heat is rejected from both sides of the radiator panels. Because of this, the effective area for rejecting heat is double the physical area of the radiator panels. The radiator design is described by Siamidis and Mason (2006).

TABLE 2.—POWER SYSTEM PARAMETERS.

	MTV		CTV
Reactor system mass	18088	kg	4973	kg
Brayton power conversion system mass	8748	kg	4374	kg
Heat rejection system mass	33456	kg	16728	kg
PMAD system mass	20484	kg	9648	kg
Radiator area (effective)	5444	m2	2722	m2
Radiator area (physical)	2722	m2	1361	m2

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Propulsion System

     This study used magnetoplasmadynamic (MPD) thrusters using hydrogen propellant. Besides operating at a high specific impulse (I_SP), the MPD thrusters also have the added advantages of a highpower capability and a compact size. This analysis used high power MPD thrusters operating at 2.5 MW_e per thruster at a constant I_SP of 5,000 sec with a thruster lifetime of 7500 hr.

     The MPD thrusters use cryogenically-stored liquid Hydrogen (LH₂) propellant. This mission utilized the 2.5 MW_e thrusters assumed in the 2002 HOPE study. The MTV vehicle used four operating thrusters for a total power level of 10 MW_e and had 4 non-operating spares for redundancy. Likewise, the CTV used two operating thrusters at a total power level of 5 MW_e and had two nonoperating spares for redundancy. The mass of the thrusters is I_SP dependant. Since a constant I_SP was used in this analysis, the mass of the thrusters was calculated using the system alpha (mass/kW_e) for an ISP of 5000 sec. The thrusters were run at an I_SP of 5000 sec due to higher efficiencies at this specific impulse. See figure 4 for the dependency of system alpha and MPD thruster efficiency versus operating I_SP.

     One power processing unit (PPU) and one radiator are assumed per thruster. The system alpha of the PPU and radiator is assumed to be 2.5 kg/kW_e. This included the mass for the power conditioning at the turbine (transformer to increase the voltage to 1 kV), the 1 kV transmission line, the PPU to convert power at the other end, and the Parasitic Load radiator (PLR) to reject waste heat. The sink temperature is assumed to be 250 K at Earth orbit for a worst-case sizing. The effective radiator areas for the two vehicles were: CTV = 273.4 m² for a rejection of 5 MW_e, and MTV = 546.8 m² for a rejection of 10 MW_e.

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Trajectory

     Both the MTV and the CTV begin in LEO at 1000 km altitude, spiral out from the Earth, and follow a low-thrust trajectory to Mars. The CTV captures into a 24.65-hr period Mars orbit (radius of periapse = 3643 km, radius of apoapse = 37,186 km, orbital period = one Martian day). The MTV spiral-captures into the same Mars orbit 12 days later. After a 60-day stay at Mars, the MTV returns the crew to Earth on a flyby trajectory. The crew aerobrakes at Earth in the ECRV for a direct-entry that is constrained to 13 km/sec at 125 km altitude. The trajectories for the vehicles were optimized using the VARITOP trajectory optimization program. The trajectory for the MTV and CTV are shown in figure 5. The dates and times of each mission phase as well as the total mission time and total crew time (for the MTV) are shown in table 3.

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     The study requirements on the total round-trip time and stay time at Mars were not optimal for an NEP mission and led to a close approach to the sun (0.41 AU) in the return trajectory. Such a close approach to the sun could require additional shielding to protect the crew, the power system, and the electronics; however, these effects were not included in this study. VARITOP was used to determine how the trip time affected the closest approach to the Sun for varying propulsion specific masses. While the propulsion specific mass does not have much of an effect, figure 6 shows that increasing the trip time actually decreases the closest distance to the Sun. Reducing the trip time could be accomplished by increasing power and/or reducing I_SP, however, this would raise the power or propulsion system masses. These trade-offs were not performed as part of this study. Since the MTV passes within the orbit of Venus (0.7233 AU), a Venus flyby may be able to improve the trip time and/or the closest approach to the Sun. A preliminary investigation using a Venus flyby, subject to the RASC mission constraints, did not show any improvement to the trip time. Similarly, another mission profile using a stay time at Mars longer than 60 days might allow for a more favorable return trajectory, but this was beyond the scope of this study, since the RASC mission requirements did not allow for changing the Mars stay-time.

Mission Mass Summary

     The mass breakdown for the MTV and CTV is shown in table 4. A common contingency factor of 25 percent was applied to all dry masses other than the RASC payloads in this analysis in order to get a final mass with contingency. A 1 percent contingency was applied to the LH₂ propellant masses.

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Conclusion

     This study has developed a space transportation architecture based on high-power nuclear electric propulsion using Brayton power conversion and magnetoplasmadynamic propulsion to support a manned Mars mission. The architecture consists of a cargo transfer vehicle with one 5 MWe fission reactor and a Mars transfer vehicle with two 5 MW_e fission reactors. The RASC study fixed the mission parameters to investigate the performance of 4 different technological approaches to accomplishing the mission. Unfortunately, these mission parameters were not optimal for this architecture and led to several difficulties. The Mars Transfer Vehicle makes a very close approach to the sun (0.41 AU) during the return trajectory in order to rendezvous with Earth. Adjusting the stay-time at Mars and/or utilizing a Venus flyby could be used to increase this distance. The trajectory sequence requires the MTV to begin thrusting before the CTV; however, this occurs before the crew is on board. The MTV similarly departs Earth and begins the heliocentric portion of the flight before the CTV, even though the CTV arrives at Mars, first. A better approach would be to launch the CTV on an earlier opportunity than the MTV to predeploy the cargo in the proper orbit before any resources associated with the MTV are launched.

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      Her room was at the bottom of the carousel, midway between ladders three and four. She followed Gaby around the curving floor, then pursued her up the ladder.

     Each rung was a little easier than the last until, at the hub, they were weightless. They pushed off from the slowly rotating ring and drifted down the central corridor to the science module. SCIMOD in NASA-ese. It was kept dark to make the instruments easier to read, and was as colourful as the inside of a jukebox. Cirocco liked it. Green lights blinked and banks of television screens hissed white noise through confetti clouds of snow. Eugene Springfield and the Polo sisters floated around the central holo tank. Their faces were bathed in the red glow.

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     Ringmaster was an elongated structure consisting of two main sections joined by a hollow tube three meters in diameter and a hundred meters long. Structural strength for the tube was provided by three composite girders on the outside, each of which transmitted the thrust of one engine to the life system balanced on top of the tube (meaning there are three engines, one for each girder).

     At the far end were the engines and a cluster of detachable fuel tanks, hidden from sight by the broad plate of the radiation shield which ringed the central tube like the rat guard on the mooring line of an ocean-going freighter. The other side of that shield was an unhealthy place to be. (radiation means fission or fusion engine. Or antimatter but that seems unlikely)

     On the other end of the tube was the life system, consisting of the science module, the control module, and the carousel.

     Control (CONMOD) was at the extreme front end, a cone-shaped protuberance rising from the big coffee can that was SCIMOD. It had the only windows on the ship, more for tradition than practicality.

     The Science Module was almost hidden behind a thicket of instrumentation. The high-gain antenna rose above it all, perched on the end of a long stalk and trained on Earth. There were two radar dishes and five telescopes, including Gaby’s 120-centimeter Newtonian.

     Just behind it was the carousel: a fat, white flywheel. It rotated slowly around the rest of the ship, with four spokes leading up from the rim.

     Strapped to the central stem were other items, including the hydroponics cylinders and the several components of the lander (the Satellite Excursion Module or SEM): life system, tug engine, two descent stages and the ascent engine.

     The lander had been intended for exploring the Saturn moons, in particular Iapetus and Rhea. After Titan—which had an atmosphere and was therefore unsuited for exploration this trip—Iapetus was the most interesting body in the neighborhood. Until the 1980’s, it had been significantly brighter in one hemisphere, but it had changed over a twenty-year period until its albedo was nearly uniform. Two troughs in the graph of luminosity now occurred at opposite points on its orbit. The lander had been designed to discover what caused it. (Novel takes place approximately in the year 2055)

     Now that trip had been scrapped in the face of the much more compelling object called Themis (the space colony).

     Ringmaster resembled another spaceship: the fictional Discovery, the Jupiter probe from the classic movie 2001: A Space Odyssey. It was not surprising that it should. Both ships had been designed from similar parameters, though one sailed only on celluloid. Cirocco was EVA to remove the last of the solar reflection panels which wrapped the life system of Ringmaster. The problem in a space vehicle is usually one of disposing of excess heat, but they were now far enough from the sun that it paid to soak up what they could get.

The deep space ship above (click on the image for full sized view) was inspired by the Travel Planner spreadsheet in the previous post, and modeled in the wonderfully simple and handy DoGA 3D modeler. The shuttle alongside is a rough approximation of the NASA shuttle, and thus a thorough anacronism in this image, but provided as a scale reference.

Of course you want some specifications of the ship. Even if you don't, you get them anyway:

Length Overall	300 meters
Departure Mass	10,000 tons
Propellant Load H2	5000 tons
Drive Mass	2000 tons
Keel and Tankage	1000 tons
Gross Payload	2000 tons
Flyway Cost	$5 billion (equivalent)

The payload includes a hab with berthing space for 50-200 passengers and crew, depending on mission duration, and a pair of detachable pods for 500 tons of express cargo, plus service bays and the like.

What this ship can do depends on its drive engine performance. If the drive puts out 2 gigawatts of thrust power — my baseline for a Realistic [TM] nuke electric drive — the ship can reach Mars in three months, give or take. (The sim gave 92 days for a 0.8 AU trip in flat space.) With a later generation drive putting out 20 gigawatts it can reach Mars in a little over a month, or Saturn in eight months.

The general arrangement of this ship is driven by design consideration — a nuclear drive that needs to be a long way from the crew, with large radiators to shed its waste heat; tanks for bulky liquid hydrogen; and a spinning hab section. Most serious proposals for deep space craft in the last 50 years have had more or less this arrangement — the movie 2001 left off the radiator fins, because in those days the audience would have been puzzled that a deep space ship had 'wings.'

A large, long-mission military craft, such as a laser star, might not look much different overall — replace the cargo pods with a laser installation and side-mounted main mirror, and perhaps a couple of smaller mirrors on rotating 'turret' mounts. Discussions here have persuaded me that heavy armor is of little use against the most likely threats facing such a ship.

Within these broad constraints, however, spaceships offer a great deal of design freedom, more than most terrestrial vehicles. Ships, planes, and faster land vehicles are all governed by fluid dynamics, and even movable shipyard cranes must conform to a 1-g gravity field. A spaceship, unless built for aerobraking, will never encounter fluid flow, and the forces exerted by high specific impulse drives — even torch level drives — are relatively gentle.

This ship might have had two propellant tanks, or half a dozen, instead of four. And the entire industrial assemblage of tanks and girders might be concealed, partly or entirely, within a 'hull' of sheeting no thicker than foil, protecting tanks and equipment from shifting heat exposure due to sunlight and shadow. Much of the ISS keel girder has a covering of some sort — in close-ups it looks a lot like canvas — that in more distant views gives the impression of a solid structure.

In fact the visual image of the ISS is dominated by its solar wings and radiators. The hab structure is fairly inconspicuous by comparison, like the hull of a sailing ship under full sail. This would be true to an extreme of solar electric ships; a 1-gigawatt solar electric drive would need a few square kilometers of solar wings. Even nuclear drives, fission or fusion, require extensive radiators — probably more than I showed — with other ship systems needing their own radiators, at varied operating temperatures. Unless the ship has an onboard reactor it must also have solar collectors for use when the drive is shut down.

All of which may do more to catch the eye than heavier but smaller structures such as the hab or even propellant tankage. And then there is color: the gold foil of the main ISS solar wings, for example.

Hollywood knows nothing of this (though I'm surprised they haven't picked up on the gold foil). Hollywood is no more interested in what real spaceships look like than it is in how they maneuver. This is only natural, even though we hard SF geeks complain. Hollywood doesn't care because its audience has almost no clue of what spaceships look like, or act like, getting most of their impressions from Hollywood itself.

The one actual spacecraft to have iconic visual status, the Shuttle, essentially looks like an airplane. The ISS has not yet acquired iconic status, though it may, especially after the Shuttle is retired. And perhaps it looks so unlike terrestrial vehicles that our eye does not yet know quite what to make of it.

As a point of comparison, watch aviation scenes in old movies, especially from before World War II. You'll see airplanes whooshing past (sometimes in pretty unconvincing special effects shots), but you will rarely see what is now a standard shot — a plane filmed from another plane in formation, hanging 'motionless' on the screen, clouds and distant landscape rolling slowly past, until perhaps the plane banks and turns away.

It is a standard shot because it is so very effective. But older movies rarely used it, because audiences would have had no idea what they were seeing. Everyone knew that airplanes were fast, and had at least some idea that their speed is what kept them in the air. A plane apparently hanging in midair would make no sense.

What changed all this, I would guess, is World War II. A flood of newsreel footage included many formation shots, and audiences gradually absorbed a feeling for what midair footage really looks like. When a postwar Jimmy Stewart enlisted for Strategic Air Command (1955), Hollywood — and its audience — were ready to see the B-36 and B-47 showcased in all their glory, including airborne formation shots.

I know what you bloodthirsty people are thinking — one good space war, and everyone will grok the visual language of space travel. Shame on you. Given enough civil space development, and time, people will get the hang of it.

The beauty of spaceships is in the eye of the beholder. The familiar aesthetics of terrestrial vehicles are as irrelevant to them as to Gothic cathedrals (which in some broad philosophical sense are themselves spaceships of a sort). General principles of design will provide some guidance. Even in making the quick thrown-together model above I found that slight changes in proportion could make the difference between a jumble of parts and a unity.

But the real visual impact of spaceships is something we will only learn from experience, by the glint of a distant sun.

Payload
Crew	25
Hab Module	100 tons
Consumables	25 tons
Other Payload	75 tons
Total Payload	200 tons
Propulsion Bus
Engine+Radiator	200 tons
Tankages+Keel	100 tons
Stats
Dry Mass	475 tons
Loaded Mass	500 tons
Propellant Mass	500 tons
Wet Mass	1000 tons

The discussion thread about 'Industrial Scale of Space' veered, among other things, into a discussion of patrol missions in space. My first reaction was that (so long as you aren't dealing with an interstellar setting) there is no place in space for wartime patrol missions. But the matter might be more complicated, and for story purposes probably should be.

According to The Free Dictionary, patrol is The act of moving about an area especially by an authorized and trained person or group, for purposes of observation, inspection, or security. This fits my own sense of the word, and is in fact a bit broader, 'security' including SSBN patrols, which are not observing or inspecting anything, just waiting for a launch order if it comes.

In a reductionist way you could say that all military spacecraft are on patrol, since they are all on orbit, and if they are orbiting a planet they have a very regular 'patrol area.' But this is not what most of us have in mind. We picture a patrol making a sweep through an area, looking for anything unusual, ready to engage any enemy they encounter, or report it and shadow it if they cannot engage it.

Back in the rocketpunk era it was plausible that, say, Earth might send a patrol past Ceres to see if the Martians had established a secret base there. But (alas!) telescopes 'patrolling' from Earth orbit can easily observe the large scale logistics traffic involved in establishing a base; watch it depart Mars and track it to Ceres. If you want a closer look you can send a robotic spy probe. If you engage in 'reconnaissance in force' by attacking Ceres, that is a task force, not a patrol.

In an all out interplanetary war there may be plenty of uncertainty on both sides, but very little of it can be resolved by sending out patrols.

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But of course all-out war is not the context in which the Space Patrol became familiar. I associate it with Heinlein's Patrol; apparently the 1950s TV series had an independent origin (unlike Tom Corbett, who was Heinlein's unacknowledged literary child).

The rocketpunk-era Patrol, which in turn gave us Starfleet, was placed in the distinctly midcentury future setting of a Federation. This is as zeerust as monorails. But plausible patrolling is not confined to Federation settings. It can justified in practically any situation but all out war.

Orbital patrol in Earth orbital space will surely be the first space patrol, and could be imagined in this century. It might initially be a general emergency response force, because travel times in Earth orbital space are short enough for classical rescue missions. On the interplanetary scale, with travel times of weeks or more likely months, rescue is rarely possible. But eventually power players will want some kind of police presence or flag showing in deep space.

As so often in these discussions, I picture a complex and ambiguous environment in which policing, diplomacy, and sometimes low level conflict blur together. To take again our Earth-Mars-Ceres example, there are kinds of reconnaissance that cannot be carried out by robots (short of high level AIs). If Ceres closes its airlocks to liberty parties from a visiting Earth patrol ship, that conveys some important intelligence information.

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The ships that perform these missions will be fairly large (and expensive). They must carry a hab pod providing prolonged life support for a significant crew: at least a commander and staff, SWAT team of espatiers, and some support for both.

Let us say a crew of 25—which is cutting the human presence very fine. Now we can venture a mass estimate. Allow 100 tons for the hab compartment plus 25 tons for crew and stores plus 75 tons other payload, for a total payload of 200 tons. Let the drive bus be 200 tons for the drive, including radiators, and 100 tons for tankage, keel, and sundry equipment.

Our patrol ship with a crew of 25 thus has a dry mass of 475 tons, mass fully equipped 500 tons, plus 500 tons propellant for a full load departure mass of 1000 tons. Cost by my usual general rule is equivalent to $500 million, perhaps $1 billion after milspecking, expensive compared to military planes, cheaper than major naval combatants.

This is no small ship. If the propellant is liquid hydrogen the tanks have a volume of about 7000 cubic meters, equivalent to a 7000 ton submarine. The payload section is about two thirds the mass of the ISS and of roughly comparable size, though the hab is probably spun giving the prolonged missions.

Armament is necessarily modest. The 75 tons of additional payload allowance probably must include a ferry craft for the espatiers and an escort gunship or two, plus their service pod, leaving perhaps 15-20 tons each for kinetics and a laser installation. The laser might be good for 20 megawatts beam power, with plug power from the 200 megawatt drive engine.

This ship is no laser star, but the laser is respectable. Assuming a modest 5 meter main mirror and a near IR wavelength of 1000 nanometers, at a range of 1000 km it can burn through Super Nano Carbon Stuff at rather more than 1 centimeter of per second. Its armament is also rather 'balanced.' My model shows that this laser can just defeat a wave of about 1000 target seekers, each with a mass of 20 kg, closing at 10 km/s—thus a total mass of 20 tons, comparable to its kinetics payload allowance.

Deploying troops, or personnel in general, is impressively expensive: About three fourths of the payload and cost of a billion dollar ship goes to support and equip a crew of 25, with perhaps a dozen espatiers. For comparison the USS Makin Island (LHD-8) displaces 41,000 tons full load, carries a crew of 1200 plus 1700 Marines, and costs about $1.8. So by my model it costs about as much to deploy one espatier as 80 marines.

And this ship is about the minimum patrol package, so standing interplanetary patrol is a costly and somewhat granular business, something not everyone can afford.

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Ray McVay Rocketpunk Patrol Ship
Dry Mass	76.2 metric tons
Wet Mass	384.6 metric tons
Mass Ratio	5
Length Z	73 meters
Length Y	20.1 meters
Length X	15.2 meters
Engine	x2 F-26-A LH/LOX
Thrust	7.7×106 N
Acceleration	0.5 g
ΔV	8,200 m/s

This is the same one from the other day, only dressed up with a nice logo and some stats. These are realistic capabilities made courtesy of the charts and other information available from Atomic Rocket and inspiration from Rick Robinson's Rocketpunk Manifesto.

My PL differs from the one in Rick Robinson's article in a few key areas. The main difference is that it is not made for long hauls. It only has a delta v of about 8200 m/s. This will not get one far in the solar system but it allows a forward deployed Patrol Craft a sufficient "range" to perform many of the missions we discussed in the last post on Building a Space Navy. Our little A-Class has enough Delta V to shape a light-second orbit around a convoy in deep space, conduct SAR missions anywhere in cis-lunar space, or to reach any moon of Saturn from any other moon. Obviously, this rocket is mostly propellant (mass ratio 5). If you drew lines through the side view of the rocket that bracket the docking rings, you would encompass the entire pressurized volume. I've got to say, it's nice to work on a warship for a change — I don't have to make it economical to run!

One of the interesting things about this design is actually the freedom the little carried craft gives me. It was a throw-away touch, originally — a design borrowed from another project. But as I got to looking at the little thing, I realized that it's about the size of the Saturn V stage/Apollo/LM stack. That means it should be able to go from Earth Departure to Lunar orbit. That means that it has the Delta V to ferry crew to and from a Patrol Craft on station away from the convoy. That means, like submarines, our Patrol Craft can have two crews and stay out for a lot longer than otherwise. This is one of those realistic touches that I hope add to the charm of the rocket's design.

ed note: a 1500 nanometer near infrared laser with a 10 meter fixed mirror can have a 4 centimeter spot size out to 220 kilometers or so. A 4 meter mirror can have a 4 centimeter spot size out to 87 kilometers or so.

Solar-electric deep space drive engines, according to Isaac Kuo at sfconsim-l, may soon achieve a power output density of about 400 watts per kilogram, when operating near Earth distance from the Sun. If you do not see what this sort of technical information could possibly have to do with so lovely an image as gossamer wings, you probably reached this blog by accident, have no poetry in you, or both.

What makes it potentially relevant as well as beautiful is that 400 watts/kg is in hailing distance of the 1 kW/kg that Isaac and I independently chose as a benchmark for nuclear-electric drive, and generally as needed for relatively fast interplanetary travel. A spacecraft using solar electric drive can thus reach the same interplanetary speeds as its cousin, though it will take somewhat longer to reach cruising speed, and somewhat longer to slow down. It is a fair prospect that with a few decades' further progress, by the time we're actually building interplanetary ships the performance of the two drives will be comparable.

This is a big deal, because solar-electric space drive is technically and operationally elegant, while nuclear-anything drive, and especially nuclear-electric drive, is not. A solar electric drive has almost no moving parts. A nuclear-electric drive has lots of complex internal plumbing to draw energy from the reactor and incidentally keep it from melting. This plumbing operates under very nasty conditions, radioactivity being nothing to sheer high temperatures.

Plumbing is a big part of what makes spaceships so expensive, because it is complicated, full of parts that can jam, and as there is never a plumber around when you need one, it has to work perfectly for months at a time. (Even if you have a plumber in the crew, taking a nuclear reactor apart en route is a pain.) Robinson's Second Law: For each gram of physics handwavium in futuristic space tech, expect about a ton of plumbing handwavium.

Nuclear drives are also full of nasty fissionable stuff, tricky and dangerous to work with, requiring heavy shielding to get anywhere near (and radiation goes a long ways in space), requiring extreme security measures in handling and storage, and socially uncomfortable no matter how careful your procedures are.

In short, anything that gets rid of nuclear reactors in space is a huge plus on every level of operation, from spacecraft construction and maintenance to obtaining funding. Solar electric drive with comparable performance banishes nuclear reactors from the inner Solar System. You don't need them for travel, and you certainly don't need them for anything else, because one thing the inner Solar System has an ample and endless supply of is sunshine. Those skies are never cloudy all day.

Solar electric power does gasp for air, or for sunshine, as you move outward from the Sun. At Mars, thrust is about half as much as near Earth. In the asteroid belt it is about a fifth to a tenth, at Jupiter one twenty-fifth, at Saturn one percent. To give this some context, a one-milligee drive, baseline performance near Earth, nudges a ship along at about 1 km/s per day, reaching orbital transfer speeds in a week or two. At Jupiter, the drive delivers some 40 microgees, and a ship puts on about 1 km/s per month, thus the better part of a year for orbital transfer burns.

The time lost due to sluggish acceleration is only half as much, some six months, and a Jupiter mission would likely be upwards of a year each way even for a nuke-electric ship. So until we have regular bus service to Jupiter, the time cost is not dreadful. The inner Solar System, through the asteroid belt, can be efficiently traveled by solar-electric drive, which ought to hold us through this century and into the next.

Of course nuclear-electric ships can be built, but Isaac also pointed out a subtle effect that could sideline them. Over the decades to come we will build solar-electric probes, and later ships, steadily developing the technology, while nuke-electric remains a paper tech, falling further and further behind. A serious advance into the outer system will require a faster drive in any case—by that time perhaps a fusion drive, which can still be two orders of magnitude below the magical performance level of a 'torch.'

Let's mentally sketch-design a solar electric ship. Departure mass with full propellant load is 400 tons, broken down as follows:

• Payload, 100 tons
• Structures and fitting, 50 tons
• Drive engine, 100 tons
• Propellant, 150 tons

The drive engine we make an advanced one, meeting the baseline standard of 1 kW/kg. Thus rated drive power is 100 megawatts. If the exhaust velocity is 50 km/s (specific impulse ~5000 seconds), 80 grams of propellant is shot out the back each second. Thrust is 4000 Newtons, about 1000 lbs, giving our ship the intended 1 milligee acceleration at full load. Mass ratio is 1.6, so total ship delta v available on departure is 23.5 km/s, enough for a pretty fast orbit to Mars.

We could 'overload' this ship with a much bigger payload, another 400 tons (thus 500 tons total payload). Max acceleration falls to half a milligee, and mission delta v to 10 km/s—still ample for the Hohmann trip to Mars, for slow freight service. Since we want to go there ourselves, we will stick with the faster version and configure it as a passenger ship. Each passenger/crewmember requires cabin space, fittings, life support equipment, provisions and supplies for the trip, plus the mass of the passenger and baggage—in all, say, about 3 tons per person, so our ship carries some 30-35 passengers and crew.

The cabin structure of this ship might be about the size of a 747 fuselage, divided into berthing compartments or roomettes, diner/lounge area, galley, storage spaces, and life support plant. If the propellant is hydrogen, the tankage will be about the same size; if other stuff is used, the tankage will be smaller. All in all, the hull portion of our ship is comparable in size and mass to a jumbo jet. As space liners go this is a modest-sized one, as its modest passenger/crew capacity shows.

Now, finally, the gossamer wings part. We accounted for the mass of the drive engine, including solar collectors, but have not yet looked at the physical size of the solar panals. They are big. Big. If we assume that about 35 percent of the sunlight that hits them is converted into thrust power, they capture some 500 watts per square meter at 1 AU—meaning that for a 100 megawatt drive you need 200,000 square meters of solar panels, a fifth of a square kilometer.

This trim little interplanetary liner is physically enormous, or at least its solar wings are. The 'wingspan' might well be one kilometer, 'wing chord' then being 200 meters. In sheer size our ship is much bigger than any vehicle ever built (though freight trains can be up to about 2 km long).

Angular, squared-off, an instrument of technology—but how can this ship be anything but a thing of beauty, an immense gleaming-black butterfly? If that is too fluttery, say a dragonfly, or to be prosaic an equally immense gleaming-black kite. Indeed the prototype configuration is much like a box kite, likely for later versions as well.

Something is magical about such ships and travel aboard them. The drive thrust and power performance is the same as for a nuke-thermal ship, but now the milligee acceleration feels appropriately gentle, not merely weak, as our ship glides from world to world on its great sun-wings. (This is not, however, solar sailing, but a sun-powered 'steamship.')

The modest capacity of this immense little ship adds to the charm. With only about 35 passengers and crew this is no tawdry impersonal cruise ship. It all has somewhat the flavor of airship travel as we imagine it—perhaps encouraged by the zeppelin-like proportions of the vehicle, the gondola dwarfed by the feather-light structure that carries it. In early decades the ship will be much more utilitarian, a transport rather than a liner—don't ring for the steward; it's your turn in the galley. But if we go to the planets we will eventually go in liners.

The scenery out the viewports* won't change much after the first week or so spiraling out from Earth. (In fact you probably ride a connecting bus up through the Van Allen belts.) By then it is time for reading, cards, conversation, and flirting, till Mars looms close and the ship begins its long graceful swoop down to parking orbit.

Bon voyage!

* I disagree with Winch. All but the most utilitarian spaceships will have a few viewports, because while there is often nothing to see, when there is it is breathtaking. And fundamentally, why else are we going into space?

• 

• 

Ce principe inventé par James R. Powell fut exploré à la même époque que NERVA par le Laboratoire national de Brookhaven. Les particules de 500 à 700 µm de diamètre étaient composées d'un noyau de carbure d'uranium UC₂ enrobé de carbone poreux (rétention des produits de fission), de carbone pyrolytique et d'une couche anti-corrosion en carbure de zirconium ZrC. Il fut proposé deux conceptions du réacteur : le réacteur à lit fixe FBR, dans lequel les particules sont stockées entre deux frittes poreuses cylindriques, et le réacteur à lit rotatif RBR qui ne possède pas de fritte intérieure (chaude) et maintient les particules contre la fritte extérieure (froide) par centrifugation à ~1000 tr/min.

Comme le RBR n'a pas de fritte chaude, il est affranchi des problèmes liés à cette pièce et peut produire une température de sortie supérieure. De plus, le moteur peut être purgé en fin de fonctionnement puis rechargé plus tard, cette possibilité évite un échauffement prolongé du moteur après son extinction (dû à la décomposition des produits de fission instables) et permet la maintenance plus aisée du réacteur. Les performances envisagées étaient une Isp de 1000 s et une poussée de 90 kN pour une masse de 1370 kg. De nombreux aspects mécaniques restèrent non résolus.

Du fait de l'important rapport surface/volume des particules, les systèmes FBR et RBR étaient réputés opérer un excellent transfert de chaleur avec l'hydrogène et annoncés capables d'atteindre des températures de 3000 à 3750 K et une Isp de 1000 à 1300 s. La zone d'échange étant très courte, un tel système a une grande densité énergétique lui autorisant une configuration plus compacte que NERVA et atteignant donc un meilleur rapport poids/poussée.

Les études de ces systèmes n'étaient pas très avancées quand elles furent stoppées en 1973 en même temps que les autres programmes de propulsion nucléaire.

---

(ed note: from Google Translate)

This principle invented by James R. Powell was explored at the same time as NERVA by the Brookhaven National Laboratory. Particles of 500 to 700 µm in diameter consisted of uranium carbide UC₂ coated porous carbon (retention of fission products), pyrolytic carbon and a carbide corrosion protective layer zirconium ZrC . It resulted in two designs of the reactor: the fixed bed reactor FBR, wherein the particles are stored between two sintered porous cylinders, and the rotating bed reactor RBR which does not have any inner frit (hot) and holds the particles against the exterior frit (cold) by centrifugation at ~ 1000 r / min.

As RBR has no hot frit layer, it is free from the problems with this and can produce higher output temperature. In addition, the engine can be purged at the end of operation and reloaded later (i.e., open the bottom and dump all the radioactive uranium dust into space), this option avoids a residual heating to the engine after throttle off (due to the decomposition of unstable fission products in the dust) and allows for easier maintenance of the reactor. The performances were considered an Isp of 1000 seconds and a thrust of 90 kN for a mass of 1370 kg. Many mechanical aspects remained unresolved.

Due to the large surface / volume ratio of the particles, the FBR and RBR systems were deemed to make an excellent heat transfer with hydrogen and announced able to reach temperatures from 3000 to 3750 K and Isp of 1000 to 1300 s . The exchange zone is very short, such a system has a high energy density to allow a more compact configuration than NERVA and thus reaching a better thrust to weight ratio.

Studies of these systems were not very advanced when they were stopped in 1973 along with other nuclear propulsion programs (Michel Van points out there were further studies from 1980s until 1992, funded by the US Department of Defense).

I was looking at the new "Scorpion" from the Skylon team last night, and trying to wrap my head around where it's supposed to get its performance and what exactly the benefit of the Helium loop is.

As far as I can tell, the main benefit — and how it gets its I_sp — is by divorcing the temperature of the hydrogen exhaust from the temperature required of the core, with the arcjet acting as a "superheater".

Extracting power from the helium to drive the electrical generator to power the arcjet may be less efficient in terms of how much of the thermal energy of the core ends up in the exhaust vs in the radiators (and note this one actually needs radiators, unlike a typical NTR), but the approach seems to be to just accept a larger reactor and wash away the pain of the added mass with a 40% higher I_sp.

• 

• 

There wasn't much doubt on the last few trips, but when we first put Slingshot together out of the wreckage of two salvaged ships, every time we boosted out there'd been a good chance we'd never set down again. There's a lot that can go wrong in the Belt, and not many ships to rescue you.

---

I shrugged and began securing the ship. There wasn't much to do. The big work is shutting down the main engines, and we'd done that a long way out from Jefferson (asteroid colony). You don't run an ion engine toward an inhabited rock if you care about your customers.

---

The entryway is a big compartment. It's filled with nearly everything you can think of: dresses, art objects, gadgets and gizmos, spare parts for air bottles, sewing machines, and anything else Janet or I think we can sell in the way-stops we make with Slingshot. Janet calls it the "boutique," and she's been pretty clever about what she buys. It makes a profit, but like everything we do, just barely.

(Nine tons of beef) I donated half a ton for the Jefferson city hall people to throw a feed with. The rest went for about thirty francs a kilo. That would just about pay for the deuterium I burned up coming to Jefferson.

(ed note: approimately 16,000 francs per metric ton of deuterium)

---

"I don't think you understand. You have half a million tons to boost up to what, five, six kilometers a second?" I took out my pocket calculator. "Sixteen tons of deuterium and eleven thousand reaction mass. That's a bloody big load. The fuel feed system's got to be built. It's not something I can just strap on and push off—"

---

I switched the comm system to Record. "Agamemnon, this is cargo tug Slingshot. I have your Mayday. Intercept is possible, but I cannot carry sufficient fuel and mass to decelerate your ship. I must vampire your dee and mass, I say again, we must transfer your fuel and reaction mass to my ship.

"We have no facilities for taking your passengers aboard. We will attempt to take your ship in tow and decelerate using your deuterium and reaction mass. Our engines are modified General Electric Model five-niner ion-fusion. Preparations for coming to your assistance are under way. Suggest your crew begin preparations for fuel transfer. Over."

---

The Register didn't give anywhere near enough data about Agamemnon. I could see from the recognition pix that she carried her reaction mass in strap-ons alongside the main hull, rather than in detachable pods right forward the way Slinger does. That meant we might have to transfer the whole lot before we could start deceleration.

---

The refinery crew had built up fuel pods for Slinger before, so they knew what I needed, but they'd never made one that had to stand up to a full fifth of a gee. A couple of centimeters is hefty acceleration when you boost big cargo, but we'd have to go out at a hundred times that.

---

They launched the big fuel pod with strap-on solids, just enough thrust to get it away from the rock so I could catch it and lock on. We had hours to spare, and I took my time matching velocities. Then Hal and I went outside to make sure everything was connected right.

---

Slingshot is basically a strongly built hollow tube with engines at one end and clamps at the other. The cabins are rings around the outside of the tube. We also carry some deuterium and reaction mass strapped on to the main hull, but for big jobs there's not nearly enough room there. Instead, we build a special fuel pod that straps onto the bow. The reaction mass can be lowered through the central tube when we're boosting.

Boost cargo goes on forward of the fuel pod. This time we didn't have any going out, but when we caught up to Agamemnon she'd ride there, no different from any other cargo capsule. That was the plan, anyway. Taking another ship in tow isn't precisely common out here.

Everything matched up. Deuterium lines, and the elevator system for handling the mass and getting it into the boiling pots aft; it all fit.

---

Ship's engines are complicated things. First you take deuterium pellets and zap them with a big laser. The dee fuses to helium. Now you've got far too much hot gas at far too high a temperature, so it goes into an MHD system that cools it and turns the energy into electricity.

Some of that powers the lasers to zap more dee. The rest powers the ion drive system. Take a metal, preferably something with a low boiling point like cesium, but since that's rare out here cadmium generally has to do. Boil it to a vapor. Put the vapor through ionizing screens that you keep charged with power from the fusion system.

Squirt the charged vapor through more charged plates to accelerate it, and you've got a drive. You've also got a charge on your ship, so you need an electron gun to get rid of that.

There are only about nine hundred things to go wrong with the system. Superconductors for the magnetic fields and charge plates: those take cryogenic systems, and those have auxiliary systems to keep them going. Nothing's simple, and nothing's small, so out of Slingshot's sixteen hundred metric tons, well over a thousand tons is engine.

Now you know why there aren't any space yachts flitting around out here. Slinger's one of the smallest ships in commission, and she's bloody big. If Jan and I hadn't happened to hit lucky by being the only possible buyers for a couple of wrecks, and hadn't had friends at Barclay's who thought we might make a go of it, we'd never have owned our own ship.

When I tell people about the engines, they don't ask what we do aboard Slinger when we're on long passages, but they're only partly right. You can't do anything to an engine while it's on. It either works or it doesn't, and all you have to do with it is see it gets fed.

It's when the damned things are shut down that the work starts, and that takes so much time that you make sure you've done everything else in the ship when you can't work on the engines. There's a lot of maintenance, as you might guess when you think that we've got to make everything we need, from air to zweiback. Living in a ship makes you appreciate planets.

Space operations go smooth, or generally they don't go at all.

---

When we were fifty kilometers behind, I cut the engines to minimum power. I didn't dare shut them down entirely. The fusion power system has no difficulty with restarts, but the ion screens are fouled if they're cooled. Unless they're cleaned or replaced we can lose as much as half our thrust—and we were going to need every dyne.

---

Agamemnon didn't look much like Slingshot. We'd closed to a quarter of a klick, and steadily drew ahead of her; when we were past her, we'd turn over and decelerate, dropping behind so that we could do the whole cycle over again.

Some features were the same, of course. The engines were not much larger than Slingshot's and looked much the same, a big cylinder covered over with tankage and coils, acceleration outports at the aft end. A smaller tube ran from the engines forward, but you couldn't see all of it because big rounded reaction mass canisters covered part of it.

---

Finally it was finished, and we could start maximum boost: a whole ten centimeters, about a hundredth of a gee. That may not sound like much, but think of the mass involved. Slinger's sixteen hundred tons were nothing, but there was Agamemnon too.

• 

• 

 Now that Elon Musk has released engineering targets for the proposed interplanetary transport system (formerly BFR), there is some meat to work with when looking at possible applications. I'm going to extrapolate, extend and abuse those numbers as thoroughly as I can after the jump.


Booster

 First off, let's look at the booster. 275 tons of dry mass, return to launch pad recovery using 7% of the fuel load (which is 469 tons) and total propellant capacity of 6700 tons. The return fuel provides about 3.5km/s of dV, allowing for the boost-back, deceleration and landing burns. All of the dV numbers to follow assume sea-level Isp for the first stage engines, so this is a conservative value. In reality the craft spends most of its time out of the bulk of the atmosphere and Isp increases rapidly into the 370's during the ascent.

 Elon mentioned that it can probably SSTO; that's true but you have to cut the landing propellant by quite a bit to make it happen. I think if the booster can survive orbital reentry then a lot of the boost-back burn is eliminated because you can simply wait in orbit until your path lines up with the launch site and let the atmosphere do most of the work. At any rate, assuming an Isp of 361 and reserving 200 tons of propellant for landing, the booster can SSTO 40 tons of payload with about 9.2km/s of dV. It would have to be something robust, or remember to subtract fairing mass from that number. Reusability after orbital re-entry for the booster is questionable, but if it works then this machine would put everyone else out of business.

 A more normal mission profile is to carry either a lander or a tanker. That's 2400 tons (lander with 300t cargo) or 2590 tons (tanker with 380t fuel as payload); in this configuration the booster provides 3.8 to 4.0 km/s of dV to the upper stage. Actual separation velocity is 2.4 km/s, so drag and gravity losses are in the 1.4 to 1.6 km/s range. (Nearly all of these losses are spent by the first stage.)

 A tanker launch is heavier, so stage 1 gives a bit less velocity (3732 m/s gross). The tanker itself has a little more juice available (6016m/s), so the stack still has about 9750m/s. The lander stack should be 3870m/s from the booster and 5625m/s from the lander, total of 9494m/s. The extra fuel in the tanker allows for orbital maneuvers. Both upper stages retain 1560-1580m/s worth of landing propellant, which is 50 tons for the tanker and 85 tons for the lander.

 The booster is expected to cost about $230 million to fabricate and be reused for about a thousand launches. Maintenance runs about $200k per launch.

Tanker

 The tanker is a lightweight 90-ton fuel tank with engines and heatshield. Fully loaded it is 2590 tons and it can deliver 380 tons of fuel to a waiting lander in orbit. It returns to the surface and lands on legs; most of the dV for this is provided by aerodynamic drag on the heatshield. The tanker is expected to cost about $130 million to fabricate and be reused for about a hundred launches. Maintenance runs about $500k per launch.

Lander

 The lander is a robust 150-ton interplanetary vehicle with 200 kW of solar panels. It can carry 300 tons to LEO. Once refueled it can carry up to 450 tons to the surface of Mars from LEO. The lander's abort to surface fuel plus five tanker trips will fully fuel the lander; Elon mentioned Mars trips with as few as three tanker trips, so not every launch window or cargo will require the full load of propellant. The pressurized volume has enough space for 100 passengers. The ship has a huge amount of extra delta-v available, so trips are expected to be as short as 90 days (90 to 150, average 115). The lander is expected to cost about $200 million to fabricate and be reused twelve times round-trip. (If it was used in LEO only then it would have reuse comparable to the tanker, around 100 flights.) Maintenance runs about $10 million per Mars flight, probably a tenth of that or less for LEO only.

Reference Plan

 In the reference plan, the lander is launched to LEO with passengers and cargo. The same booster launches a tanker three to five times; the tanker docks with the lander, transfers fuel, lands, reloads and repeats. Once the lander is ready to go it departs during the transfer window. Passengers cruise for about 115 days in microgravity using currently-available life support tech. The lander performs an aerocapture in Mars atmosphere with direct descent, flying sideways during the hot parts for maximum drag and then landing propulsively on the tail. ISRU equipment makes propellants for the return trip. When it is time to return, the lander launches from Mars surface to low orbit and shortly after departs for Earth. Earth arrival is just like Mars: aerocapture into direct descent. Passengers and cargo unload, then the ship gets a deep maintenance overhaul.

 The first few flights carry a small ISRU plant as part of their cargo. This is enough to produce return fuel during one transfer window using the ship's solar array and probably some extra panels. Components would be an atmosphere compressor, ice excavators and water extraction oven for the raw materials, then electrolysis and Sabatier process equipment to make propellants, followed by liquefaction equipment to turn them into liquids. The very first flight will be unmanned and possibly the one after; passengers won't be sent until there is return propellant available. Later flights will rely on a built-up ISRU facility for refueling, freeing up some cargo capacity.

 Musk expects the cost for these flights to eventually drop below $140k per ton of payload to Mars surface. If components don't meet their re-use targets the cost would go up, but even if the lander only gets used twice the price is still around $300k per ton. None of this includes the ~$10 billion estimated development costs. Consider that he had Raptor engine test firing video and pictures of a 12-meter carbon fiber propellant tank ready for the IAC this year, plus they've been doing simulations and refinements for about a decade; this is going to happen and it's going to happen soon. The only data they don't have in enough resolution is Mars EDL for larger objects with supersonic retropropulsion and a map of accessible water ice; the upcoming Red Dragon missions will help fill in the gaps.

Going beyond the plan

 Elon mentioned several interesting destinations throughout the solar system, up to and including Kuiper belt objects provided there are propellant depots available. The easiest targets would be those with atmosphere for aerocapture, and anything with possible ISRU would be a good target as well. Targets beyond the main belt will likely require nuclear power of some kind.

 A lander with no payload and ~85 tons of landing fuel has about 8.2km/s in the tank. The landing reserve is about 1.5km/s, so if you run the tanks dry you can get about 9.9 km/s. That's enough to go from Earth to nearly any main belt object (Ceres, Pallas, etc.) and make orbit. Payload to Ceres orbit would be 19 tons, for example, or 5 tons to Ceres surface. If you launch from Mars orbit fully fueled then you can just barely orbit Pluto on a Hohmann transfer (~46 years travel time), or you could take 70 tons to the surface of Vesta and back.

 The tanker is a better option for deep space. It's lighter and carries more fuel, so in an expendable configuration it has about 12.5km/s available. At $130 million that's pretty affordable for a deep-space probe bus, especially a chemical one with over 12km/s dV.

 I'm working on getting a trajectory optimization tool running with current data, so I haven't had a chance to find reasonable dV numbers for missions to Titan or other gas giant moons. It would be a pretty wild ride to aerocapture through Saturn or Jupiter and land on one of the moons, but if the mission is launched from Mars orbit you could put down quite a bit of payload. Hoping to have more numbers to work with soon.

So, this Tuesday SpaceX pulled back the curtain to announce their Interplanetary Transport System—a monstrously large rocket, fully reusable and about two and a half times the size of a Saturn V moon rocket—capable of transporting a hundred people to Mars, and with a goal of initial flight testing within a decade.

It's not total vaporware: in the past couple of weeks they also tested the first full-up Raptor engine that will power the ITS (a cryogenic methalox engine with a closed-cycle gas generator, which gives it a specific impulse head and shoulders higher than Apollo-era kit and the capability to operate on fuel generated from the Martian atmosphere for return flights). They've also unveiled the biggest carbon fiber tank ever assembled (the fully-reusable ITS will use carbon composites extensively), and have unveiled a bunch of targets for what the ITS stack will be able to achieve: in non-reusable form it will be able to deliver a 500 tonne payload to LEO, and with reusability in mind a 320 tonne interplanetary craft capable of landing vertically on Mars (and, when refuelled, of returning to Mars orbit without staging).

So, here's my question:

What are the other possible commercial applications of the ITS, besides sending a million optimists to Mars?

Here's what I can see:

0. 1-2 order of magnitude cost reduction in cost/ton of payload to orbit: this is axiomatic. ITS won't be commercially viable for Musk's proposed Mars colonization bid if the per-launch cost of this big-ass fully reusable rocket significantly exceeds that of the big-ass but not fully reusable (the second stage is disposable) Falcon Heavy that flies later this year. So let's posit a cap of $100M on flight costs, or maybe $400M for a disposable shot (which would only really be necessary for a single monolithic payload that can't be broken down into sub-elements massing less than 300 tons—candidates for which, see below). (Here are SpaceX's cost estimates.)

1. Big, dumb, comsats: Currently the mass of a geosynchronous comsat is constrained by the payload of the available boosters, which are tailored to fit the perceived requirements of the comsat market. About half the mass of a comsat in GEO is fuel, used for positioning (satellites in geosynchronous orbit drift, very gradually, away from their parking longitude). Their power output is constrained by the solar panels they can carry and the size of their emitters. So a big GEO comsat today is on the order of 5-8 tons. A current advanced geosynchronous comsat such as Inmarsat-4A F4 has a 12 kW electrical system; this obviously puts a ceiling on its broadcast power; but ITS raises the bar so high that it effectively disappears. The first post-ITS generation of comsats could have power outputs in the megawatt range if necessary. So I'm going to guess that 1-2 decades after ITS flies, we're going to see satellite phones converge with regular cellphones in terms of size, convenience, and bandwidth capacity (although they're going to cost more). Upshot: terrestrial 5G and hypothetical 6G high bandwidth service will look more like municipal-area gigabit wifi, and your phone will cut over to satellite bandwidth if you roam into rural areas (or even suburban areas, by the US definition). But you won't notice anything except a slight increase in latency. It's as if your cell tower just moved into orbit.

2. No more Kessler syndrome nightmares: the launch stack is fully reusable. Anyone not aiming to operate a reusable launch stack by 2030 at this point is a buggy whip manufacturer. So that's one source of debris gone. And another source of the problem is the number of objects in space. A few giant satellites are less likely to shed debris or risk a collision hazard than a large number of small satellites. And we'll have so much spare lift capacity that cleanup becomes a practical possibility, paid for by the insurance underwriting industry: sending up a fleet of cubesats to hunt down, grapple with, and de-orbit 1960s paint chips is cheap compared to the payout if said paint chip holes your orbital Hilton.

3. Space tourism, for realz: the Bigelow BA-2100 spacehab only needs a 70-90 ton LEO launch capacity and has half the volume of the entire ISS. We can conservatively estimate that a space hotel with a ~300 ton mass fabricated using Bigelow's expandable tech and flown on the ITS would have 3-4 times the habitable space of the ISS, so room for 20-40 tourists and staff. (The inflatable hab tech isn't vapourware either: there's one docked to the ISS right now.) A week in space won't be a cheap vacation, but Virgin Galactic think people will pay $25K for 10 minutes in free fall; I reckon $250,000 for a honeymoon in orbit will find some takers among the 1%. (Passengers would travel as a sub-cargo aboard an ITS which would be mostly carrying other types of paying cargo.)

4. Return to the Moon, this time for good: a huge problem with proposals to build a permanent base on the Moon is that the Moon is short on volatiles that you can turn into fuel, and has no atmosphere worth mentioning for aerobraking purposes. (Lithobraking is not recommended. Or should I say lithobreaking.) One serious proposal for a long-term Lunar presence requires the construction of a Lunar space elevator. This would not run from surface to geosynchronous orbit—the moon, being tidally locked, has no GEO—but instead to the L1 (near-side) or L2 (far-side) Earth-Moon libration points, 56,000 and 67,000 kilometers from the surface (points where the effect of the Moon's gravity and the effect of the centrifugal force resulting from the elevator system's synchronous, rigid body rotation cancel each other out and an elevator could be stable). Unlike a terrestrial space elevator sufficiently high tensile strength materials for such a tether already exist. There is, however, the slight problem of fabricating and shipping a 120,000 kilometer long cable out to near-Lunar orbit (and capturing a near-Earth asteroid to act as a counterweight). This is just a wild-ass Charlie guess, but I suspect shipping up 500 tonne cable drums will work out cheaper in the end than trying to build a carbon fiber factory in space (at least, until space industries are sufficiently developed to go the whole eat-your-own-dogfood distance). (Upshot: ITS probably makes the folks at LiftPort Group very, very happy.

5. Stupidly enormous space telescopes: Because there is a budget and a booster that can lift primary mirrors 17 meters in diameter is going to make the astronomical community need a change of underwear when the implications sink in. (Put it this way: one part of the value proposition is "maps of continent-sized features on terrestrial exoplanets" by 2040.)

6. (Speculative) Wake shield molecular beam epitaxy fab lines: with a wake shield you can produce an ultra-hard vacuum suitable for growing rystalline semiconductor thin films. I don't know wht the commercial implications are other than really pure GaAs and AlGaAs semiconductor substrates, but with rock-bottom launch costs and the ever spiralling cost of semiconductor fab lines (part of which is down to the requirement for clean room air flow on a large scale) we might see some semiconductor manufacturing activities planned for deployment in orbit after 2030. (After all, high-end microprocessors—at least before they're sliced, diced, and packaged in pin grid arrays—are some of the few objects that cost so much per unit weight that they'd be worth retrieving from orbit even with current generation flight costs.)

Anyway: these are the first non-Mars short term applications of ITS that I can come up with off the top of my head. Stuff I don't think is plausible: ITS upper stage derivatives used as ballistic point to point passenger transports on Earth (because reasons), pick-axe wielding asteroid miners going out to the belt to hew mineral ore and bring it back to Earth orbit (yeah, we'll get asteroid mining, but probably by using the smallest feasible robotic gravity tractor—you don't need the ITS for that job), microgravity crystallography factories for pharmaceuticals (oh come on), Lunar 3He mining for aneutronic fusion reactors (because if you can do aneutronic fusion at all Boron is much cheaper). Anything else?

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Glossary of Acronyms

AL Airlock
AMU Astronaut maneuvering unit
BOD Bail-out Device (BOW or stranded BOR)
BOR Bail-out and Return device
BOW Bail-out and Wait device
DV Distressed Vehicle
EC/LS Environmental Control and Life Support system
EVA Extravehicular Activity
PAL Portable Airlock
SRV Space Rescue Vehicle

G.3.2 Personnel Carrier and Auxiliary Aids

     The transfer of injured personnel from the distressed vehicle to the rescue vehicle without further injury or damage can be a significant factor in assuring containment of the medical situation. Injuries requiring careful handling and immobilization include fractures and/or dislocations. Such injuries can result from moving in a weightless environment, body acceleration during maneuvering or docking operations, meteoroid penetration of the spacecraft cabin or the spacesuit during EVA, and mechanical injuries arising from explosive decompression, explosions, and walking on extraterrestrial surfaces

     The ideal characteristics of a device to transport an individual with such injury include:

1. Light weight, with minimum storage volume
2. Provision for body and limb restraints
3. Protection against bumping interior surfaces while being moved
4. Handles or grips, and tie-down provisions to the spacecraft interior

     One concept combining these characteristics visualizes a stretcher-type inflatable air mattress with bumping shields, restraint belts and hand holds, and compressed air bottle. Restraint belts would be provided for both the torso and for each leg.

     To provide full immobilization for fractures and/or dislocations, the use of pneumatic splints could supplement this personnel carrier. The storage volume of the carrier uninflated is estimated at 0.25 cubic foot (0.007 m³) with a total weight of under 10 pounds (4.5 kg).

G.3.3 Other Equipment With Medical Utility

     Medical conditions on board the DV may require means for quarantining and/or decontaminating members of the DV crew and/or members of the rescue crew. Appendix H discusses two equipment items which could have secondary application in this context. The transfer capsule was conceived as a device to allow transfer of ill or injured personnel unable to don pressure garments for EVA transfer when docking was infeasible. This capsule, equipped with an independent environmental control system, could be docked against the SRV during the return-to-haven phase while serving as a one-man quarantine station. The portable airlock could hold two men for this purpose. This airlock could also be equipped to perform biological decontamination functions for personnel transferring in a docked situation or during a quarantine period.

     These devices could also be used to isolate against radioactive contamination. In that role, docking against the SRV may not be feasible and tethering at a suitable separation distance may be required.

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H.3.2.1.2 Despin Devices

     The causes and magnitudes of uncontrolled motion of a distressed vehicle have been discussed in Appendix I. Action to reduce this uncontrolled motion can be taken by either the DV or the SRV. It is also possible to postulate a scheme for despinning which, in the event of total failure of DV command systems, and/or of the DV crew, could be activated by the SRV either remotely or by sending a crew in EVA.

     Three basic schemes for despinning were considered; the application of external torques, energy dissipation within the DV, and inertia augmentation. All three schemes lend themselves to pre-positioned devices within or on the DV; only the first and the third method could also be applied by the SRV.

     Examples of external torques are the use of reaction control systems already provided on the DV, or the application of external thrusters attached by the SRV crew. If the size relationships between SRV and DV are appropriate, grappling mechanisms on the SRV may be able to couple the two vehicles to allow the propulsive capability of the SRV to reduce the motion.

     Without provision of special equipment, energy dissipation within the DV is often available in the form of sloshing propellants or magnetic forces such as eddy forces. Such inherent dissipating processes tend to act very slowly, possibly requiring weeks to produce the desired stabilization. Special energy absorbers in the form of fluid hoops are also conceivable, which may speed up the stabilization process.

     Inertial augmentation can be provided by extendable masses on booms or weights on cables (Yo-Yo System).

     A brief analysis was performed to size two such feasible systems which also offer the possibility of being brought to the DV by an SRV and attached either manually by a crew or by a remote-controlled manipulator. For both systems, the characteristics of the uncontrolled DV motion must be known to reasonable accuracy to permit the sizing of the control forces and the proper locating of the attachment point. The mass-on-cable and the rocket thruster concepts were selected for analysis and were applied to a tumbling space station.

     The assumptions concerning the characteristics of the space station were as follows:

• Weight of the station = 120, 000 lb
• Motion around the major axis of rotation
• Rate of motion = 4 rpm

     It was also assumed that attachment aids had been provided on the station in anticipation of the need.

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     If the tumbling mode requires despin device attachment at unpredictable positions, the concept of prepositioned despin aids is not applicable. Further study of this problem is necessary prior to the selection of any de spin device.

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H.3.2.2.2 Entry to DV

     As already indicated, entry to the DV from EVA may require EVA operations to force entry hatches, the removal of modules already docked against the entry hatch, etc. The use of manipulators of either open platform or capsule type may be required in this operation. One other consideration applies in the instance when the DV hatch through which entry is to be made is not equipped with a working airlock. When entry under EVA conditions is to be made into a DV which has retained all or some of its atmosphere, and where continued retention of the atmosphere is essential, an airlock cycle must be performed either in a nominal airlock or by evacuating and repressurizing the DV compartment behind the entry hatch. If compartment pressure cycling is infeasible due to lack of functioning equipment, or due to the presence of a shirtsleeve crew (who you presumably do not want to kill by asphyxiation), a portable device may be required which can serve as an airlock. Such a portable airlock (PAL) could be of expandable design in order to reduce stowage volume requirements in the SRV and could have other additional functions. It could, for example, be utilized between docked spacecraft to serve as an atmospheric contamination barrier between DV and SRV. Equipped with appropriate chemical spray systems, it would also prevent biological contamination of the SRV, if the DV emergency has created such a hazard. Used as a BOD or as a quarantine device it would require more extensive EC/LS provisions,

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     A conceptual arrangement of a PAL sized for two astronauts is shown collapsed for stowage in Fig. H-20. The flexible center section, made of material that can be folded, is extended by pressurization to a length long enough to accommodate a suited astronaut in a stretched-out position.

     The PAL consists of two active ring-and-cone assemblies, an extendible cylindrical member, a cylindrical structure which encloses the collapsed flexible member, and a breathing and pressurization subsystem. The two active ring-and-cone assemblies incorporate the docking mechanism and the hatches and are connected by the folded flexible cylindrical member. The airlock thus permits entry into the DV by an astronaut operating in an EVA mode or by direct transfer to the DV from a rescue vehicle docked to the opposite end of the portable airlock. A typical flexible material having the required structural and packaging properties is the Goodyear "Airmat." The PAL is extended initially by using the pressurization system which also provides the breathing atmosphere.

     The docking hatches combined into the docking mechanism at each end of the portable airlock are identical in size and provide a clear 5.0 ft (1.5 m) diameter opening for transfer of equipment. The two docking mechanisms are fastened together in the stowed position by the rigid cylindrical structural member which encloses the collapsed flexible member. The rigid cylindrical member incorporates a circumferential joint, located midway along its length, which is held together by spring loaded locks which are released either electromechanically or by the internal pressure used to extend the airlock into the operating position. The rigid cylindrical member also provides protection for the extensible material during stowage. Possible methods for retracting the airlock after use include telescopic tubes, cable retraction devices, extendible booms, etc. The stowed volume of the airlock is about 380 ft³ (10.8 m³) and its weight is estimated at about 1600 lb (725.7 kg).

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H.3.2.2.3 Exit from DV

     Much of what has already been discussed under transit and entry into the DV will, of course, also apply to the exit phase of the rescue mission. A PAL is as necessary to exit as to entry unless the rescue crew has been able to provide every member of the DV crew with a pressure garment, thus permitting the decompression of the DV compartment prior to exit. Pressure suits are also required if the vehicles are not docked. However, many medical situations can be postulated for a crew disabled by the emergency which may prevent dressing at least some of the DV crew in pressure suits.

     Broken arms and legs are examples of such situations. In such an instance, the concept of a transfer capsule might be valuable. Such a device would also be stowed in the collapsed condition within the SRV in order to reduce storage volume requirements.

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     A capsule design concept for transferring men and equipment between the rescue vehicle and the DV is shown in Figure H-21. A North American Rockwell hatch design, featuring a hatch within a hatch, was selected as a representative design. The 5.0 ft (1.5 m) outer diameter hatch corresponds to the transfer tunnel diameter used in the space station design. The inner auxiliary hatch is approximately 3.0 ft (0.9 m) in diameter. This hatch is large enough to permit passage of a personnel carrier defined in Appendix G for transporting an injured astronaut. This inner hatch is also large enough for transporting emergency equipment into the crew transfer capsule. Modifying the North American Rockwell docking hatch to include a latch ring permits attaching the crew transfer capsule directly to the hatch, thus eliminating additional docking fixtures. This concept results in a smaller diameter attachment and reduced weight.

     The transfer capsule consists of two major components, an inflatable member and a cylindrical metal shell structure approximately 36.0 inches long attached to the inflatable member. The part of the shell structure that attaches to to the DV hatch latch ring is designed to incorporate a number of docking latches located radially around the shell. These docking latches engage the inside lip of the latch ring and achieve attachment to the hatch in a manner similar to that described for the attachable docking fixture. An inflatable pressure seal is provided between the capsule and the hatch. The cylindrical metal shell structure contains a removal hatch that is mounted approximately midway inside the shell. The hatch is removable in a manner similar to the Gemini heat shield hatch. The inflatable section is inflated to a shape similar to that shown in the sketch by pressurizing the capsule with breathing atmosphere provided from high pressure storage containers.

     In the stowed position, the inflatable portion of the capsule is folded and packed inside the metal shell portion of the capsule. The stowed volume and weight are estimated at 50 ft³ (1.4 m³) and 500 lb (227 kg).

     After the astronaut has been placed into the capsule, the hatches are resealed and the capsule is transported to the SRV by manipulators or by the rescue crew with AMUs. Attached to the SRV, the astronaut may be removed from the capsule or may be restricted to the capsule for a quarantine period, with life support provided from the SRV.

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H.3.2.3.1 Docking Interface

     A brief conceptual analysis was undertaken to determine whether SRV's could be equipped with soft docking fixtures capable of reducing the difficulty of docking to a DV with some residual wobble. The analysis was non-quantitative; stress analysis was not performed and the design was not matched to specific values of DV motion.

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     The soft docking fixture shown in Figure H-22 is configured to accommodate slight motions between the rescue vehicle and the DV. If the DV motions are greater than can be accommodated by the docking fixture, these motions must be reduced to a tolerable level. The concept calls for flexibly mounting the North American Rockwell docking design with a neuter docking device and a passive ring. The docking port on the DV is assumed to be a passive ring assembly. This concept could be modified into a ring/cone assembly which can be mated with another active ring/cone docking assembly to form a complete neuter docking subassembly.

     Further study of this concept is required to derive methods for extending the flexible bellows toward the DV shell to provide a pressure seal and the correct stiffness at the flexible connection to minimize vehicle dynamic interactions resulting from differential vehicle motion. The weight increment of this type of docking fixture over the conventional design would be about 250 lb.

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     A damaged spacecraft implies the possibility of a situation where docking facilities are unavailable. If a space station is taken as an example, many of its docking ports will be occupied by experiment modules. Other ports may have logistic vehicles such as space tugs or the EOS docked to them. Finally, the emergency situation calling for rescue may have destroyed some of the ports or may have closed the passage between them and the space station compartment which the rescue crew is attempting to reach. EVA airlocks may be provided on the station but may not have been equipped with docking fixtures. The SRV may thus be faced with the necessity of creating an opening against which it could dock. In either case, that of an opening already available such as an EVA air lock, or that of an opening that must be cut into the hull, a docking fixture must somehow be placed over the opening to permit SRV docking. The concept of such a portable docking fixture was briefly investigated.

     The portable, attachable docking fixture shown in Figure H-23 permits docking to a distressed vehicle via an EVA port. For purposes of this study, a 6-ft diameter opening, 12 inches larger than the standard hatch opening, was assumed. This larger opening permitted the use of the North American Rockwell docking design with minor modifications, and also permitted the use of ramp-shaped docking pawls identical in cross-sectional shape to the North American Rockwell docking cone. Space is also available for a standard 5 ft diameter hatch for transfer of personnel and cargo. The portable docking fixture is secured within the 6 ft diameter opening by the eight docking pawls located radially about the opening. The pawls are engaged initially at the pawl tips: continued movement of the fixture farther from the port opening causes the docking pawls to rotate over center about the pivot points as pressure is exerted on the ramp portion of the pawl. The pawls continue to rotate about their respective pivots until the end points of the ramps have been reached; the spring-loaded pawls then snap into place behind the DV opening, thus securing the fixture between the back face of the pawl and the docking fixture seal face. An inflatable seal between the seal face and the DV port area prevents pressure loss as the DV is repressurized. Retraction of the locking devices to permit withdrawal is provided through the use of electromechanical or completely mechanical devices.

     The basic concept can also be applied to an opening specifically cut into the pressure hull of a space vehicle, providing the structure had initially been designed to permit this.

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Stuhlinger Ion Rocket
Length	150 m
Wet mass	360 metric tons
Dry mass	Lander ship: 240 metric tons Cargo ship:170 metric tons
Cesium Propellant	Lander ship: 120 metric tons Cargo ship: 190 metric tons
Mass of Mars Lander	70 metric ton
Storm cellar mass	50 metric tons
Storm cellar height	1.9 m
Storm cellar diameter	2.8
Crew	3
Rotation rate	1.3 rpm
Artificial gravity	0.14 g
Reactor thermal power	115 MWt
Generator power	40 MWe
Radiator area	4,300 m2
Radiator dissipation	75 MWt
Propulsion	Ion
thrust	98 N

ABSTRACT

     This paper presents a concept for a manned expedition to the surface of Mars, to take place in the early 1980's. By using a fleet of five electrically propelled vehicles, high probabilities of mission success and crew safety are provided, whereas individual vehicle mass and complexity are kept within limits that appear reasonable for the time period considered. The five vehicles carry three men each and are similar in mass, design, and operating characteristics. Three vehicles carry high-thrust landing craft for transportation between the orbiting electric ships and the Martian surface. The other two carry compensating amounts of propellant during the outbound portion of the interplanetary journey. A constant, slow rotation of the large vehicles is maintained throughout the journey to provide a simulated gravity environment in the crew compartments. Each crew compartment contains a heavily shielded radiation shelter. The main vehicles are powered by fission reactors, from which electric power is derived by turbo-generator, thermionic, or plasma dynamic conversion techniques. The electrostatic thrust devices operate under constant conditions but are programmed in direction during the interplanetary transfer. Details on the flight plan, as well as conceptual design data on the vehicles, are presented.

INTRODUCTION

     It is generally accepted that a manned expedition to the surface of Mars will he carried out soon after such an ambitious project becomes technically feasible. Although a number of authors have published preliminary studies of manned planetary missions, the technological resources required for their implementation are considerably beyond our reach at present. As a consequence, no formal programs leading to manned Mars landings have been established, but it seems reasonable to predict that detailed design studies on the necessary vehicles will begin toward the end of this decade. This prediction is based on the theory that the Mars expedition will be the natural follow-on project to he undertaken after the lunar program reaches enough maturity that lunar surface operations are being carried out more or less routinely. By that time, our knowledge of space radiations, meteoroid hazards, and the planets and their surroundings will have advanced far beyond its present state. In addition to the knowledge and experience gained from extensive manned operations in Earth-moon space, sophisticated probes will have gathered detailed information on the physical characteristics of the nearer planets and the environment of solar space.

     Even lunar projects involving manned landings can become quite expensive in terms of vehicle requirements and costs, as is shown by the present Manned Lunar Landing Project. The Mars project, however, represents a total propulsion requirement at least several times the magnitude of the manned lunar landing. The only means available for avoiding the excessive growth of vehicle sizes and costs in the more difficult missions is the increase of engine specific impulse through the use of advanced propulsion devices. The prime candidates in this field are the nuclear-heated and the electrical types, and, of these two, the electrical (electrostatic or ion) system promises higher specific impulses by an order of magnitude or more. A major task in realizing the potential of electric propulsion is the development of power sources having high power-to-mass ratios (specific power α). As outlined later, it is assumed in the present study that this development work will result in multimegawatt sources with α values near 0.5 kw/kg, Under this assumption, electric propulsion is the clear choice for total Mars mission durations over 400 days (1). Shorter times involve excessive growth factors with either propulsion system. The compromise between mission duration and payload fraction must be made rather arbitrarily at this time. The present authors feel that considerations of crew adaptation and equipment reliability place considerable importance on the reduction of total mission time. A rather short total mission duration of 572. days was chosen for the present study.

     For any manned mission, a high probability of safe crew return will be required. In addition, Mars landing expeditions will be quite expensive, and so economic considerations also will call for a high probability of mission success. The resulting reliability problem is aggravated by the fact that the manned Mars expedition represents such a large increase in difficulty over its lunar forerunners. The method adopted herein for dealing with these problems consists of using a small fleet of vehicles, one or more of which can be abandoned in the event of a permanently disabling failure.

PRINCIPAL ASSUMPTIONS

Mission Profile

     It is assumed that the planetary terminals for the electric vehicles will be low-energy orbits (low-altitude satellite orbits) about the Earth and Mars. This choice is made because it offers a very significant reduction in the overall mass ratio for the mission, as compared to low-thrust operation between high-energy satellite orbits. The magnitude of this saving can be estimated from the fact that the low-thrust planetocentric trajectory segments as assumed herein involve a total mass ratio of 1.16, whereas high-thrust propulsion through the same energy states (between low satellite orbits and escape) requires a single-stage mass ratio of 7.76 for the chemical rocket (at 450 sec I_sp) and 2.57 for the nuclear rocket (at 1000 sec I_sp). Also, since orbital assembly is indicated by the size of the electric vehicles, this full utilization of the electric propulsion system simplifies the orbital operations procedures. The use of high-energy terminal orbits would involve either assembly problems aggravated by the less accessible "site" or the assembly of a larger, more complex vehicle (high-thrust stage attached) in a low orbit. The foregoing advantages appear to outweigh the disadvantages, chief among which are longer total mission time (25 to 30%) and increased time spent in the Van Allen belts. However, since heavy crew shielding is needed for solar flare protection in any case, the additional weight required for Van Allen shielding is not prohibitive.

     Another important assurnption regarding the mission profile has to do with the interplanetary trajectories. The most economical choice from the propulsion standpoint would he direct transfers on both the inbound and outbound legs. but this profile requires waiting times at Mars which are considered excessive (e.g., 455 days using two Hohmann 260-day transfers). By assuming an indirect trajectory (passing inside Earth's orbit) for the return leg, the waiting time at Mars is reduced to 29 days, although the interplanetary return leg now requires 120 days more than the outbound leg. The general form of the complete mission profile is shown in Fig. 1.

Vehicle Allocations

     As mentioned previously, it is assumed that a fleet of several vehicles will be used for the Mars expedition. There are two main advantages in this approach: in case failures occur, one or more vehicles can be abandoned without aborting the mission or losing crew members; and individual vehicle masses are reduced by sharing the transportation job between vehicles. The following paragraphs outline the reasoning behind the vehicle allocations adopted.

     To reduce the landing craft mass, two are used, one carrying equipment and supplies and the other carrying the landing crew. Under this plan, an alternate return craft is available on Mars before the manned landing craft is dispatched. A third craft is held in hack-up status, to be used in the event one of the first two fails. Each landing craft, fully fueled, will have a mass of 70 tons. Assuming one landing craft per electric vehicle, three of the craft-carrying vehicles are required. However, since it appears desirable that the expedition be able to tolerate the loss of more than one main vehicle, two additional electric vehicles are allocated. These allow a further reduction in individual vehicle mass, since they can carry increased propellant loadings instead of landing craft.

     Thus the fleet consists of five electric vehicles, three (designated Type A) carrying landing craft, and the other two (Type B) carrying compensating amounts of propellant. The propellant loadings are equalized between vehicles after the landing craft are detached for the return trip. Each vehicle is manned by a crew of three. It carries their life support supplies and equipment, including a radiation shelter. The shelters are large enough to accommodate additional crew members in case a vehicle must be abandoned; one shelter could even hold the entire mission complement of 15 under crowded conditions during an emergency return to Earth.

Power Supply

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     The Prime source of power will be a fast fission reactor, developing a total amount of heat power of the order of 115 MW. Located at one end of the oblong vehicle (Fig. 2), the reactor will he shielded by a relatively small shadow shield. The coolant from the reactor will transmit the heat energy through a heat exchanger to the working fluid of the conversion system. No attempt was made in the present study to specify the nature of the electric generator, except that the overall efficiency of the energy transformation from heat to electricity is about 35% and that the specific power or α, referred to the entire thrust-producing system, is 0.5 kw-kg^-1. Several generator systems appear feasible at megawatt power levels: turbo-electric generators, thermionic generators, and plasma dynamic generators. At the present time, the turbo-electric system is farthest advanced, although no power supply of this kind has been flight-tested so far. Thermionic and plasma dynamic generators are showing promise in laboratory tests. They have the advantage that no moving parts are necessary in their operation. However, extensive engineering development will be necessary before one of these systems can be committed to a large space vehicle project.

     The determining factor in the efficiency of a turboelectric power generator is the upper temperature limit to which the turbine can be exposed. Great improvements in high temperature materials were achaieved in recent years, but the operation of a turbine at 1450°K inlet temperature over a period of almost two years is still a formidable problem. The generator, too, would have to work under elevated temperature because of its proximity to the turbine. The area of the radiation cooler determines the condenser temperature. For minimurn system mass at specified output power and turbine inlet temperature, the condenser temperature should be 3/4 of the inlet temperature, provided that the conversion efficiency is 100% (ideal Carnot cycle) and that the mass of the prime power source is proportional to its heat power output. In more realistic systems, the ratio of condenser to inlet temperature should be 2/3 to 1/2 (2).

     The present study has a parametric nature only. An inlet temperature of 1450°K and a condenser temperature of 750°K were assumed, At an estimated overall efiiciency of 35%, 75 out of the original 115 MW must be rejected as waste heat; this situation requires a total radiator surface of approxirnately 4300 m². The radiator will consist of a large number of elements, as shown in Fig. 3. ln case of meteoroid damage, the damaged element will be shut off; the radiator is sufficiently overdesigned to allow the anticipated number of damaged elements to be closed off.

     A specific power of 0.5 kw-kg^-1 (4.4 lb-kw^-1) is still beyond our present technologies. However, realistic design studies are available today which promise a specific power of 0.2 to 0.3 kw-kg^-1 in the 300 kw to 1 Mw range (3). On that basis, it appears justified to expect a specific power of 0. 5 kw-kg^-1 for a 40 Mw system in the 1975 to 1985 period Hopefully, the increasing demands of the space program will assure that strong emphasis is placed on power source development. It is probable also that progress in other important programs, involving research in such fields as plasma physics and high-temperature materials, will yield advances that are applicable to the power source effort.

THE MASTER PLAN

     In line with the foregoing discussion, it appears reasonable to assume that a fairly elaborate manned expedition to Mars will be undertaken in the normal course of space flight development, perhaps in the early 1980's. The use of electrically propelled vehicles for the main part of the journey is also indicated, with high-thrust carriers required for transportation between planetary surfaces and satellite orbits. Based on these assumptions and the vehicle allocation selected previously, a master plan for the Mars expedition can be formulated. The following outline lists the major steps in sequence:

1. Inject modular sub-assemblies into low satellite orbit about Earth. Assemble five vehicles, check out, and place flight crew aboard.
2. Execute Earth escape maneuver, travel direct trajectory to Mars, and descend to low satellite orbit.
3. Dispatch landing craft with equipment.
4. Dispatch landing craft with several crew members.
5. Spend 29 days exploring, making measurements, and collecting samples.
6. Return manned landing craft to electric vehicle, transfer crew and nonexpendable cargo, abandon landing craft, and equalize propellants between ships.
7. Escape from Mars, travel indirect trajectory to Earth, and descend to low satellite orbit.
8. Pick up crew and other payload with Earth-based carrier, leaving electric vehicles in orbit for re-use.

     In the event a vehicle fails during the Earth-to-Mars journey and has to be abandoned, its crew could be transferred to the other vehicles, allowing the mission to continue Conceivably two vehicles could fail without aborting the mission, provided the two are not of the same type. Otherwise the expedition would return to Earth as promptly as possible. The most practical return trajectory would either be computed on-board or be radioed to the vehicle from an Earth-based station.

     If the first landing craft (carrying equipment) were lost during the Mars landing operation, the second would he dispatched unmanned with a substitute cargo. The manned back-up craft would be launched only if the substitute cargo craft lands successfully, thus providing an alternate return craft on Mars.

CREW SHIELDING

     The problem of providing adequate radiation shielding for crew members during an extended interplanetary journey is poorly defined at present. Our knowledge of penetrating solar radiation, its character, spectral distribution, and time variation is rather uncertain. The situation is somewhat better regarding radiation from the outer Van Allen belt, and this source evidently requires much less shielding than the inner belt. The possibility that Mars is surrounded by trapped radiation is recognized, but there is no basis at present for making any assumptions. Finally, there is considerable disagreement on the question of what radiation dose should be considered tolerable for crewmen on extended space flights. The authors believe that the accepted dose will be closer to 100R (1 Sievert) than to the much lower working limits now being applied in industrial exposures. This view is based. partly on the assumption that greater risks will be deemed appropriate in such "occupations" and partly on the expectation that, as biological effects knowledge improves, a decreasing need for conservatism will be reflected in dose limit specifications.

(ed note: This was sort of the case. Current astronaut limits are 1.5 Sieverts/month, 3.0 Sieverts/year, and 4.0 Sieverts career limit)

     In view of the foregoing uncertainties, very approximate methods have been used in designing radiation shelters for the vehicles now being considered. As shown in Fig. 2 (inset), the shelter consists of a graphite shell (90 g/cm²) (understand that 90g cm^-2 is the same as 90 g/cm²), around which are placed layers of other materials carried on-board primarily for other essential purposes (oxygen cylinders, propellant, drinking water, and miscellaneous supplies and equipment). In addition to the shelter itself, a thin shield of approximately 3 g-cm^-2 encloses the entire crew compartment. This external shield consists of a 1.5 g/cm² outer layer of low-Z material and a high-Z inner layer of equal weight for bremsstrahlung attenuation.

     Most of the propellant will have been exhausted during the return trip, but this shielding loss is partially compensated by stacking the samples from Mars around the shelter. Also, since the Van Allen belts are traversed in about half the outbound passage time, the shielding loss due to propellant depletion will he acceptable.

     The internal volume of the shelter is approximated by a cylinder 2.8 m in diameter and 1.9 m high. This space accommodates the usual three crew members quite comfortably, but it can hold the entire mission complement of 15 if an emergency arises requiring return to Earth in a single vehicle. The mass of material carried specifically for shielding is approximately 50 tons. The total amount of shielding around the shelter is about 190 g-cm^-2 during the entire outbound trip. During the inbound trip, it will gradually decrease from about 180 g-cm^-2 to 130 g-cm^-2. The dose rates to be expected inside the shelter are very difficult to estimate. Assuming a crossing time of 20 days through the inner Van Allen belt on the outbound, and of 10 days on the inbound trip, the accumulated dose may be on the order of 60 to 80 R. Representative figures for solar flare doses may be as much as about 1 R for low energy solar proton events and about 10 R for high energy events (4). Low energy solar proton events occur at a rate of about 10/year during solar maximum, whereas high energy flares are more than an order of magnitude less frequent.

TRAJECTORY AND PROPULSION

     In the present analysis, it is assumed somewhat conservatively that the electric vehicles operate under a constant thrust program; i.e., the thrust at any time is either a selected discrete value or zero. This mode of operation has the advantage that the propulsion system is required to operate under only one set of conditions, simplifying the design of the power supply and the ion motors. It has been shown (e.g., Ref. 5) that better performance can theoretically he obtained through the use of a variable thrust propulsion system, but the development of such a system is still far in the future, In the event that variable thrust propulsion does become practicable before the Mars mission is undertaken, a somewhat higher payload fraction will be attainable.

     The electric vehicle flight path as conceived here consists of two general types of trajectories: planetocentric spirals and interplanetary transfers. The spirals occur in regions where the planetary gravitational fields are dominant, and the propulsion problem is primarily one of changing vehicle energy (achieving planetary escape from a satellite orbit, or the reverse process). A constant tangential thrust program is chosen here for the spiral maneuvers, since it has the advantage of simplicity and near-optimum performance. Numerical data for the specific spirals employed were obtained from the generalized constant tangential thrust trajectory charts of Moeckel (6).

     The choice and definition of propulsion programs for the interplanetary transfer segments is considerably more difficult, in that the velocity vector of the target planet must be matched by the vehicle at the time of arrival. This additional requirement can he met by providing the vehicle with the capability of thrust direction programming. A number of optimized transfer trajectories for such vehicles have been computed recently by Melbourne, MacKay, and others (5, 7, 8). On the basis of these and other references, the interplanetary trajectory and propulsion parameters chosen here were found to be compatible, although all the trajectory figures have not yet been fully verified by direct detailed computation. The propellant mass allowances will be more than adequate.

     Besides producing the required transfers between the planetary orbits, the transfer trajectories must be timed so that rendezvous with the target planet is achieved, This constraint can he specified rather simply under the assumption of circular, coplanar planetary orbits. For the Mars rendezvous, the only requirement is that the vehicle leave Earth at the proper time in the synodic cycle of Mars (as determined by the transfer time and angle of the trajectory selected). The condition for Earth rendezvous is that the heliocentric angle subtended by the complete trajectory, excluding the geocentric spirals, be equal to the angle swept out by Earth during the corresponding time interval. This condition can be re-stated:

θ = ω_Eτ

where ω_E is the mean motion of Earth, τ is the time spent away from Earth's vicinity, and θ is the angular displacement of the vehicle during τ

Table 1: Flight Plan

Mission Phase	Duration of phase (days)	Elapsed time at end of phase (days)
Escape Earth	56	56
Earth-Mars transfer	148	204
Descent to low orbit about Mars	21	225
Wait in orbit	29	254
Mars escape	18	272
Mars-Earth transfer	268	540
Descent to a low orbit about Earth	32	572

     In the present case, τ is 484 days (Table 1). Since ω_E is about 0. 986 deg/day, the value of θ required to produce rendezvous at Earth is 477°. The outbound and inbound transfer angles are estimated to be 120° and 320°, respectively. Since Mars passes through an angle of nearly 36° while the vehicle is bound to it (68 days), the Earth rendezvous condition is approximately satisfied. The Mars rendezvous condition is ignored in the present case, since it involves only the specific calendar phasing of the mission.

     Having determined the approximate propulsion requirements for the mission, it is desirable to apply some type of optimization procedure in the design of the vehicle. Specifically, decisions must be reached regarding the operating characteristics of ion motors and the mass fractions to be allocated to propellant and to the propulsion system. The procedure applied here is described in Ref. 9. It consists essentially of determining the exhaust velocity level at which the propellant and propulsion system masses are apportioned so as to maximize payload mass. The values for these quantities shown in Table 2 are the result of this type of optimization process. Table 2 also gives other derived quantities and summarizes the major characteristics of the vehicles.

Table 2: Vehicle specifications

Mass summary (tons)
	Type A (3 veh.)	Type B (2 veh.)	Return (5 veh.)
Propulsion system	80	80	80
Radiation shelter	50	50	50
Landing craft	70
Propellant	120	190	88
Samples from Mars			12
Net payload	40	40	40
Total	360	360	270
Operating characteristics
Total electric power		40,000 kw
Beam power		36,200 kw
Specific power (thrust-producing system)		0.5 kw-kg-1
Exhaust velocity		140 km-sec-1
Mass flow rate		3.7 g-sec-1
Voltage		14,000 v
Current		2,700 a
Thrust		530 Newtons
Initial acceleration		1.47 × 10-4 g

DESIGN FEATURES AND VEHICLE CONFIGURATION

     The general form of the proposed electric vehicle is shown in Fig. 2. Rotation of the vehicle about its "yaw" axis is provided for two purposes: it creates an artificial gravity environment for the crew, and it facilitates vapor-liquid phase separation in the radiation coolers. To minimize any adverse effects of Coriolis forces on the crew, the angular rate is maintained at a rather low value, about 1.3 rpm. The resulting acceleration felt by the crew is approximately 10% of standard gravity (actually more like 14%).

     Solid rockets are used to impart rotation to the vehicle initially and to compensate for spin decay later as needed (because restartable hypergolic fuels were considered too dangerous). The plane of rotation is kept as near parallel to the ecliptic plane as possible, to minimize solar radiation incident on the radiators and to minimize the rotation plane changes required during the journey. The small plane corrections that remain are accomplished by operating the two groups of ion engines independently for short periods. The vehicle's center of gravity is preserved by maintaining the proper distribution of propellant reserves in the two tank sections.

     The radiators consist of a large number of parallel tubes with flat fins extending on both sides. Although the type of the electric generator is not specified in the present study, it is assumed that the hot side of the conversion system has a temperature of about 1450°K. The radiation cooler is supposed to be of the isothermal condenser type with a temperature of about 750°K. Beryllium appears to be the most desirable radiator material because of its low density and its high strength even at elevated temperatures (3).

      The thickness of the pipe walls in the radiator is determined by the potential meteoroid damage. Based on data. reported by Whipple (10), it is expected that with a pipe wall of 8-mm thickness the radiator will be punctured about 15 times during the duration of the mission. The radiator construction proposed here includes valves to shut off those pipes that are hit, allowing radiator operation to continue through the undamaged pipes.

     A model of the manned Mars vehicle is shown in Fig. 4.

• 

---

REFERENCES

1 Moeckel, W.E., "Fast interplanetary missions with low-thrust propulsion systems," NASA TR R-T9 (1961).

2 Pitkin, E.T., "Optimum radiator temperature for space power systems," ARS J. 2.9, 596 (1959).

3 Huth, J.H., "Space vehicle power plants," Handbook of Astronautical Engineering, edited by H.H, Koelle (McGraw-Hill Book Co. Inc., New York, 1961), pp. 15-40.

4 Keller, J.W., verbal communication, NASA George C. Marshall Space Flight Center (1961).

5 Melbourne, W.G. and Sauer. C.G., "Optimum thrust programs for power-limited propulsion systems," Jet Propulsion Lab. TR. 32-118 (June 15, 1961).

6 Moeckel, W.E., “Trajectories with constant tangential thrust in central gravational fields," NASA TR R-53 (1960).

7 Melbourne, W. G., Sauer, C.G., and Richardson, D.E., "Interplanetary trajectory optimization with power-limited propulsion systems," Inst. Aerospace Sci. Paper (November 1961).

8 MacKay, J.S., Rossa, L.G., and Zimmerman, A.V., "Optimum low-acceleration trajectories for Earth-Mars transfer," Inst. Aerospace Sci. Paper (November 1961).

9 Stuhlinger, E. and Seitz, R.N., "Electrostatic propulsion systems for space vehicles,“ Advances in Space Science. edited by F.I. Ordway III (Academic Press. New York, 1960), Vol. 2, pp. 263-349.

10 Whipple, F.L., "The meteoritic "risk to space vehicles," 8th International Astronautical Congress, Barcelona (Springer-Yerlag, Vienna, 1958), pp. 413-428.

One of his new duties had involved overseeing the operation of The Rock’s propulsion system. Like most large spacecraft, the asteroid was powered by antimatter. Thousands of power packs had been shipped from the big power satellites. These were simple toroidal pipes filled with hard vacuum and surrounded by self-sustaining magnetic fields. Each contained enough antimatter to power a normal spacecraft for a hundred round trips to the Moon. Yet, each fed The Rock’s massive ion engines for less than a day before exhaustion.

It had taken four years of powered flight to move The Rock into an orbit that ranged from 800,000 to 1.2 million kilometers above the Earth.

(ed note: 4 years at 1 antimatter toroid per day = 730 total antimatter toroids)

---

     Barnes was unfazed by the answer. “My bank has studied the economics of asteroid capture. They estimate it to be ten times as expensive as a similar project on Luna. Why is that?”
     “Lots of reasons,” Thorpe replied. “The Rock masses 300 billion tons. A propulsion system to move that much mass does not come cheap. Then there is the time involved in the project. It took four years, you know. That bears on the cost of money, insurance, and wages. Finally, there is the fuel cost. We ate up nearly ten kilograms of antimatter getting The Rock into Earth orbit.”

(ed note: 10 kilograms of antimatter in 730 toroids means about 14 grams of antimatter per toroid. 2.52×10¹⁵ joules per toroid, about 600 kilotons)

     “Ten kilograms, did you say? That’s quite a lot, isn’t it?”
     Thorpe nodded. “About two years’ production for one of the big power satellites.” (5 kg antimatter per year from one power satellite)
     “It was my impression that The Rock orbited quite close to the Earth. Almost hit it, in fact! Why so much antimatter?”
     “It’s true that The Rock’s initial orbit occasionally brought it quite close to us. However, it was also inclined ten degrees to the ecliptic.”
     “The what?”
     “The plane in which the Earth orbits. Change-of-plane is the most costly of all space maneuvers. Eighty percent of the antimatter we burned went to realigning The Rock’s orbital plane. After that, getting from solar to terrestrial orbit was easy.”

---

As the clock reached zero, magnetic fields were rearranged and a few nanograms of antimatter injected into the ship’s (lunar landing craft) thrust chamber. There they encountered a powerful jet of water. Antimatter encountered normal matter and combined in a burst of raw energy. The resulting temperature rise turned the water directly into plasma. Within milliseconds, an incandescent plume leapt downward from between the landing craft’s huge splayed feet and its descent began to slow.

---

     “To be blunt, Mr. Thorpe, my analysts have checked and find that your request for fuel is far in excess of what you really need. Your own mission plan calls for the expenditure of 4.5 grams of antimatter and eight million kilograms of monatomic hydrogen. Yet, you requested nine antimatter grams. That is exactly double your projections.”
     “I know that,” Thorpe growled.
     “If I may remind you, Mr. Thorpe, my department has been charged with ensuring that we obtain maximum efficiency out of our limited funds and Mr. Smith has a belt tightening campaign in progress at the moment.”
     “Tighten someone else’s belt. I need that antimatter!”
     “We feel that five grams would be more than sufficient for your needs,” Monet said.
     “That’s only ten percent above our rock bottom needs.”
     “Eleven percent. My people have checked industry practice and ten percent energy reserves are quite common.”
     “On the milk run to Luna Equatorial Station! Damn it, we are going out to chase a comet! We have to assume that things won’t go precisely as we planned them.”
     “A good manager doesn’t allow such deviations to occur, Mr. Thorpe. My staff assures me that our allocation is more than fair.”
     “Your staff isn’t risking their collective asses. I am. So is every man and woman spacing aboard Admiral Farragut. Either we get the full nine grams or we don’t leave orbit.”
     “My God, man! Do you know what antimatter closed at on today’s market?”
     “I know what my life is worth. More importantly, I know what Mr. Smith will say if he has to settle this dispute.”
     The comptroller stiffened. “You are, of course, free to take the matter to him. I doubt he will approve squandering the corporation’s funds, however.”
     Thorpe took a deep breath and decided to approach the problem from another direction. “Look, the price of antimatter is expected to rise steadily for the next couple of years, right?”
     “That is correct. Apparently, the Avalon Project is expending far more energy than was originally appreciated.”
     “Then we merely sell any excess on the open market when we get back. Even figuring the cost of money, we should be able to break even. We might even turn a profit.”

---

Ten kilometers aft-orbit lay the PowerStat itself. Like the half dozen other power stations that orbited 37,000 kilometers above the equator, Sierra Skies was a collection of intricate mechanisms flying in loose formation. The habitat cylinder was rotating slowly in the undimmed sunlight, its red-and-white checkered hull bright against the black of space. Around the habitat lay six large fusion generators, giant spheres from which emanated long towers adorned by paddle-shaped radiators. The radiators glowed white-hot. Each generator produced 1200 bevawatts of electrical power ("beva" is an obsolete metric prefix supplanted by "giga". 1200 bevawatts = 1.2×10¹² watts). Half of this output was sent along thick cables to the orbiting rectenna, where it was converted into a low-density microwave beam and transmitted down to Earth.

The powerstats created a great deal of energy that they did not transmit to Earth. This they used to synthesize antimatter. The antiprotons were manufactured in particle accelerators, then cooled, converted into antihydrogen, and stored in superconducting magnetic traps. The overall efficiency of the process was less than ten percent (which is fantastically efficient. Dr. Forward figures we will be lucky to get 0.01%). Even so, antimatter was the best power source yet devised for spacecraft. The dual nature of the powerstats’ business had long been the subject of controversy. Was it more important, the argument went, to deliver energy to Earth or to synthesize antimatter for the ships of deep space? To those who lived beyond the atmosphere, there had never been any argument. A steady supply of antimatter was as necessary to their lives as oxygen or ice.

(ed note: my slide rule says 1200 bevawatts per fusion generator divided by 2 is 6×10¹¹ watts. 10% efficiency is 6×10¹⁰ watts. Divided by 1.50327759×10^-10 Joules/antiproton, multiplied by 1.672621898×10^-27 kilograms/antiproton gives 6.7×10^-7 kilogram of antiprotons produced per second or 1 kilogram of antimatter every 1,497,925 seconds or 17.3 days. Per generator, 1 kg per 2.9 days with six generators. 126 kg per year, which conflicts with the 5 kg/year specified in the novel.)

(ed note: Working the other way, five power stations produce 25 kilograms of antimatter per year. This means each station produces 5 kg of antimatter per year, or 1 kg of antimatter every 73 days or 6,307,200 seconds. that is 1.59×10^-7 kilograms of antimatter per second. Divided by 1.672621898×10^-27 kilograms/antiproton then multiplied by 1.50327759×10^-10 Joules/antiproton gives 1.42×10¹⁰ watts. 10% efficiency swells that to 1.42×10¹¹ watts or 142 gigawatts. Divided by six fusion generators gives 24 gigawatts per generator. Which conflicts with the 1200 gigawatts per generator specified in the novel.)

(ed note: Since I cannot leave well enough alone I try to rationalize the figures by taking the 1200 bevawatts per fusion generator and 1 kg of antimatter every 73 days and figure the manufacturing efficiency is not 10% but closer to 0.39%. Which is still much better than Dr. Forward's 0.01% )

---

“I hear the same thing from the ship,” Amber said. “Luckily, chasing the comet doesn’t require the same amount of reaction mass that the run out from Earth did.”

Admiral Farragut’s original design had included only a single hexagram of spherical hydrogen tanks (4000 m³ each, 24,000 m³ total for hexagram) around the ship’s power module. That capacity had been tripled for the Comet Hastings Expedition — to a total of 72,000 cubic meters — by changing the normal spherical tanks out for cylindrical ones. The extra volume of reaction mass, plus the antimatter plasma the ship had taken on at the Sierra Skies PowerStat, gave the converted freighter the ability to change its velocity by 100 km/sec. Most of that delta V capability had been required to reach Jupiter six months after launch. To replace all the hydrogen they had used on the outbound leg would require the launch of more than one hundred fuel cylinders.

Fortunately, chasing the comet as it left the Jovian system required only that the ship accelerate to 15 km/sec above Calistan orbital speed. With another 5 km/sec added for safety, two-dozen replenishment cylinders would provide sufficient reaction mass to continue the mission. Amber had heard that there was a certain amount of wastage, which brought the number of ice containers to be launched to thirty. Once at the comet, of course, Admiral Farragut’s crew would have two years in which to refine the reaction mass they would need to return home.

---

     “Whatever’s bothering you, I’d say that you’ve figured out the energy required to accelerate a 500 kilometer ball of ice 5 centimeters per second.”
     “How did you know?”
     “I move asteroids for a living, remember? How bad is it?”
     “Bad,” she replied. “Everyone seems to forget that the nucleus masses 60 million billion tons. That’s one-twentieth the mass of all the water in all the oceans on Earth!”
     “It’s a lot,” he agreed. “So how hard will it to be to accelerate it that tiny bit?”
     “I figure the job will take 125 billion tons of reaction mass and a quarter-ton of antimatter.”
     “And that is what has you worried?”
     She sipped from her bulb, and then nodded. “You know it’s a good year when the human race manages to synthesize even 25 kilograms of antimatter (each station can produce 5 kg per year, so there must be 5 power stations). How are we going to get two hundred and fifty in the next eighteen months?”

---

For explosive power they would use antimatter bombs at the bottom of deep bore holes. Calling them “bombs” was something of a misnomer. They were actually standard antimatter storage rings to which small packets of chemical explosives had been affixed. When the explosives were detonated, the magnetic fields of the storage rings would fail, releasing a full kilogram of antimatter into the surrounding ice. The resulting energy release would be channeled into the fault system, hopefully causing a titanic steam explosion beneath several square kilometers of the asteroid’s surface. Thunderstrike ought to split open like a ripe melon, ejecting Ground Zero Crater and its surroundings into space.

---

Goliath, Gargantua, and Godzilla had been built to haul bulk material between Luna and the space habitats in the days before the Luna City mass driver. Once the mass driver was completed, its efficiency had proved too much for the big ships. They and their five sisters had been placed in orbital storage. They had remained there for nearly thirty years before the Thunderstrike Project took them over. For the past two months, five hundred skilled workers had worked round the clock to modify them for the coming mission. The old chemical engines had been ripped out and the latest antimatter powered converters installed. Inside their layers of protective magnetic fields, the antimatter chambers operated at temperatures approaching a million degrees centigrade (so instead of an exhaust velocity of 55,720 m/s it is more like 196,000). Where it had taken Admiral Farragut six months to reach Jupiter, Goliath and her sisters would make the same journey in only 12 weeks.

The winged shuttle fired its attitude jets as it approached Goliath. The bulk carrier was a large sphere 150 meters in diameter. Most of this, Barbara knew, was internal tankage for reaction mass and consumables. Much of the heavy cargo was to be carried externally, welded to hard points all over the ship. As they made their approach, Barbara saw everything from medium crawlers to large laser drill rigs arrayed around Goliath’s hull. The arrangement gave the ship a messy look, and probably played hell with the captain’s center of gravity calculations, but made off loading equipment at the nucleus relatively easy.

---

     “Captain Jacques Marché, of Godzilla. These antimatter bombs that each ship is to carry. How many of them will there be?”
     “As many as possible, Captain. At the moment, we have budgeted 38 kilograms of antimatter for the explosives. You will appreciate that we had to strip every stockpile in the system to get that much.”

With the advent of lunar exploration and round trip lunar transport, both chemical and nuclear, there inevitably will arise malfunctions and emergencies. There will arise communication difficulties, navigational errors, propulsion breakdowns, and structural failures. There are possibilities of collisions between spacecraft and of fatal damage from matter in space. More likely, however, are onboard concerns of life-support malfunctions, auxiliary power irregularities, compartment over pressurization (in some cases, explosions), cargo shifting, and unforeseen disorders. These are the realities of increased space travel.

In anticipation of spaceflight realities, there would be need for a nuclear rescue ship operating in translunar space. The primary role of such a ship would be to save human life and those extraterrestrial specimens aboard any ill-fated lunar vehicle. A secondary role would be to salvage the spacecraft if at all possible.

This means that the rescue ship would require propulsive capability to drastically change orbit planes and altitudes. It would require excess ΔV to proceed with dispatch to rendezvous with a disabled spacecraft. In addition, capability would be required for transferring personnel and equipment, making repairs to a disabled vehicle, and even taking it in tow if conditions warranted. The latest advances in crew facilitation, passenger accommodations, repair shops, navigational devices, and communication equipment would be required. As an introductory concept, one arrangement of a nuclear rescue ship is presented in Figure 11-11 (see above).

A particular feature to note in Figure 11-11 is the use of two nuclear engines. Each engine would be of the lunar ferry vintage and, therefore, would be sufficiently well developed and man-rated for rescue ship design. These engines would be indexed by a nominal I_sp of 1000 seconds; they would have a short time overrating of, perhaps 1100 seconds. This overrating implies conditional melting of nuclear fuel in the reactor for emergency maneuvers and dispatch.

A rescue ship would be characterized by a large inert weight compared to a regular transport vehicle. This means that large magnitudes of engine thrust would be required. However, during periods of non-emergencies, low thrusts could be used. The vehicle F/W_o characteristics (Thrust-to-weight ratio) would vary over a wide range: possibly from 0.1 during non-emergencies to 1 during emergencies. Two engines would provide the high thrust capacity for emergencies. During non-emergencies, one engine could be left idling; the other engine could provide low thrust for economic cruise. Furthermore, two engines would provide engine-out capability for take-home in the event of malfunction in one of the engines. For reactor control reasons, the two reactors would have to be neutronically isolated from each other. For this purpose, note the neutron isolation shield in Figure 11-11.

(ed note: Nuclear reactors are throttled by carefully controlling the amount of available neutrons within the reactor. A second reactor randomly spraying extra neutrons into the first reactor is therefore a Bad Thing. "Neutronically isolated" is a fancy way of saying "preventing uninvited neutrons from crashing the party." Related term is "Neutronic Decoupling")

• 

A suggested patrol region for the rescue ship is indicated in Figure 11-12 (see above). Note that a rendezvous orbit has been designated so that the rescue ship could replenish its propellant from the nuclear lunar transport system. By having rendezvous missions with nuclear ferry routes, rescued personnel, lunar specimens, and damaged spacecraft parts could be returned to Earth without the need for the rescue ship returning. Also, rescue ship crew members could be duty-rotated this way. This would increase the on-station time of a nuclear rescue ship.

      The Nixon was obviously a reworked USSS3 (U.S. Space Station #3). It still had three parallel tubes, side by side and spaced a hundred meters apart. The two outer tubes, each a hundred meters long, still contained the ship’s living quarters and were still known as Habitats 1 and 2. The center tube, the axle, contained storage and the shuttle bay.
     The axle, instead of stopping where it intersected the rear connecting elevator tubes, continued back for another two hundred and seventy meters. About halfway back on the extended axle—all of it still zero-gee—were the engineering section and the twin reactors of the nuclear power plant. At the aft end, another hundred meters away from the reactors, the VASIMRs were still under construction. Between the reactors and the engineering system were a cluster of spherical tanks that would hold the thousands of tons of water that would make up the reaction mass for the VASIMRs.

---

     And then there was Becca’s answer to the waste heat problem. In the nine weeks since the Chinese had launched, she’d moved heaven, Earth, and no small number of recalcitrant engineers. Between Engineering and the reactor modules, two four-hundred-meter-long masts projected out from the center axle, from Sandy’s viewpoint, one “up,” one “down.” At the end of each mast was what Sandy, who had done some sailing, thought of as a spar—but which also looked like the crossbar on a capital letter T.
     Two more beams, the same length as the top spars on the T, projected out “horizontally” for a hundred meters on both sides of the axle. They held the extrusion nozzles for the molten radiator alloy that Becca Johansson would use to cool the reactors. The “T” spars would re-collect the now-frozen alloy, effectively thin sheets of foil, and send it back to the reactor.
     Dozens of nearly invisible guy wires ran from the booms and masts to the axle. The guy wires were made of graphene composites that tied all the pieces into a rigid structure, far more inflexible and lightweight than any equivalent scaffold of metal. It reminded Sandy of an unfinished box kite, all balsa wood struts and string. At this moment, the struts were bare.

     Shortly, there would be sails.

     The engineers and design teams had fallen in line behind Becca Johansson’s scheme for handling the Nixon’s prodigious power demands. Not because they were happy with it; they just couldn’t think of anything else that would let them beat the Chinese to Saturn.
     Grumbling, they designed a ceramic core reactor that ran at a glowing-yellow temperature and heated the primary coolant—pressurized liquid sodium—to over nineteen hundred degrees Celsius. The superheated liquid sodium ran through a heat exchanger where it boiled more sodium. That vapor drove the primary ceramic composite turbines at two hundred atmospheres and nineteen hundred degrees Celsius.
     The sodium vapor condensed downstream of the turbine, in a secondary heat exchanger, where it heated steam to a supercritical eight hundred and eighty degrees. That drove the next set of turbines. Extreme as all of this was, it didn’t justify an epithet like harebrained.

     The final stage was another matter.

     Downstream of the secondary turbines, the steam, cooled to six hundred and fifty degrees Celsius, entered the heat exchanger for the ship’s radiators. It melted radiator alloy, a eutectic blend of aluminum, magnesium, and beryllium that liquefied at six hundred degrees Celsius. In doing so, it absorbed nearly two hundred watt-hours of heat per kilogram of melt. All Becca had to do was get rid of the heat in the molten alloy, and that was what merited the epithet.
     Her heat exchanger extruded the alloy into space in molten ribbons a meter wide and a tenth of a millimeter thick. Cold rollers in the extrusion nozzles clad the ribbons in a skin of frozen and roughened alloy, just microns thick.
     The rough skin improved the ribbons’ heat radiation properties and kept the thin, wide bands of liquid alloy from breaking apart into a spray of droplets. As they sped toward the spars four hundred meters away, the ribbons cooled and froze, radiating tremendous amounts of energy into space.

     It was a brutally efficient scheme for dumping the vast amounts of heat, but it was tricky.

     The nearly liquid ribbons of metal had to be guided electromagnetically as they squirted out from both sides of an extrusion boom, and then led to the spars over four hundred meters out. There, the solidified ribbon was fed by rollers mounted on the spars to the central masts and back down into the melting pot. Managing one ribbon was a technically challenging feat. For Becca’s system to get the Nixon to Saturn, it had to extrude and control hundreds of them, all at the same time.

     Thus, sails—or, for the more poetic, a moth with huge wings and a tiny body.

     Each of the sails comprised almost a hundred ribbons, running side by side from the boom that extruded them to the spars, across the spars and back to the heat exchanger reservoir. The alloy circulated perpetually, hundreds of semi-molten ribbons in constant motion, safely disposing of the reactors’ waste heat. When the ship was under full power, 150,000 square meters of the dull silvery metal—the equivalent of twenty-eight American football fields—would radiate nine gigawatts of heat into space. So said the theory.
     As for the practice, the first full-scale test would give them a good idea of what worked, and what might not.

     The power engineers had to bring the reactors online to produce enough heat and power to test out the turbines and the boilers and melt the alloy reservoir of the heat exchanger. But they couldn’t go too far, too quickly, because the relatively puny auxiliary cooling system had to handle the thermal load until the main heat exchanger was fully operational. It was a delicate matter. Reactors of this design didn’t really like being run at less than one percent of their rated output. If an instability got out of hand it could result in a core meltdown, and that would be the end of the mission, and possibly the space station.

     But after a week, life got interesting again. The reactors were as happy as they were ever going to be; the heat exchanger reservoir was stable at its operating temperature of just over six hundred degrees Celsius; and all the guidance sensors were nominal. Becca had taken a deep breath and given the instruction to open one slot nozzle, at minimum operating pressure.
     Slowly, slowly a tenth-millimeter-thick, meter-wide ribbon of metal crawled out of the boom toward one of the spars. It wavered for a moment, wobbled, and then the guidance sensors and control magnets latched onto it. Dedicated supercomputers analyzed the ribbon’s hesitant path and issued instructions to guidance magnets to induce precisely formulated eddy currents into the ribbon. Electromagnetism did its part; the ribbon was forced back onto the straight and narrow toward the waiting spar.
     After two minutes, the leading end of the ribbon reached the recovery spar, was picked up by the rollers, and fed across the spar and back down the mast. Engineering broke out in cheers. Sandy was happy; it was dramatic. That languorous silver band creeping across four hundred meters of space was great for building tension, and Sandy planned to include every second of that footage in the final cut. Make the audience sweat the same way the engineers had.
     The engineers opened the second nozzle and extruded a second meter-wide ribbon. It behaved much like the first. There were three hundred and fifty more of these to go. Allowing for pauses for status checks, the engineers would be at it for eighteen hours before all four sails were fully deployed.

     The next morning, back in his egg, Sandy watched as four giant frosted-pewter rectangles of metal, hundreds of meters in size, ran from the spars to the booms, like square-rigged sails. The alignment was so perfect that from a distance the sails looked like single sheets instead of hundreds of parallel ribbons of radiator alloy.

     Ramping up the heat exchanger-radiator system was simple enough in concept; it just required speeding up the extruders. The faster the metal got fed into space, the faster they could dump waste heat. Currently the extruders were streaming ribbons at a leisurely three meters per second, but in full operation the ribbon velocity would be over a hundred and sixty meters per second. The plan for the day would be to take the ribbons to ninety meters per second. If that worked, the system would be taken down while Sandy and the other engineers went over every bit of data produced by the dozens of recorders watching the event.
     Even at the slower ninety meters per second, everything needed to work hand in hand perfectly. The heat exchanger needed sufficient heat coming in from the reactors to keep the alloy reservoir molten. If the extruders ran too fast for the reactors, the exchanger would dump too much heat into space and the reservoir would cool down. If the temperature dropped below the six-hundred-degree melting point of the radiator alloy, the reservoir would freeze up and the engineers would have to shut it down. So the reactors depended upon the heat exchanger to keep from melting down, and the heat exchanger depended upon the reactors to keep from freezing up.
     In an hour, the reactors were up to five percent, dumping nearly a gigawatt of heat into the exchanger and radiators. Ribbons streamed out at twenty meters a second. The reactor managers sounded happy. The heat exchanger engineers sounded happy. Dr. Johansson sounded slightly less stressed than usual.

     The reactors were up to twenty percent. The heat exchanger was happily dissipating 3.5 gigawatts, its ribbons zipping along at seventy meters per second. Out of the corner of his eye, Sandy (in a space pod) caught a glimmer of light off one of the sails. Something different. He said, “I’m seeing something different out there, what’s . . . Hey! You guys! Joe! Cassie! Back up! Back up! Get out of there, get away!”
     A second later, the sails exploded. That’s what it looked like, anyway, from Sandy’s vantage point. The three hundred and fifty-two silvery metal ribbons making up the sails broke free of their lock-step, straight-as-arrow paths and went flying wildly into space, thin silver streamers spewing out in all directions like Christmas tinsel.
     No permanent damage done to the station or to the mission, no human injuries of any kind, though Johansson sounded mightily pissed off. From what he could tell, the ribbon guidance system had cratered. As the speed of the ribbons increased, an instability appeared. From what he could hear over the audio links, the engineers didn’t know if it was vibration in the extruders or some sort of feedback loop between the ribbons and the sensors and the control magnets, or if the computer controls hadn’t been up to the task.
     Whatever, those fast-flying ribbons had developed wobbles and the wobbles had grown uncontrollably until finally the whole control system collapsed under impossible demands and the ribbons started flying off in all directions.

---

     “The Nixon has four VASIMR engines, two coupled to each reactor/power subsystem. Each engine, full on, gobbles down over two and a half gigawatts of electricity. Combined, they suck up more juice than many major cities. What the Nixon gets for all that juice is thrust. For those of you with scientific minds, at launch, the VASIMRs will deliver over two hundred thousand newtons of thrust. That sounds like a lot, except each of the Chinese’s ten nuclear thermal engines (nuclear lightbulb) produces five times as much thrust as Nixon’s entire complement.
     “The Nixon is not a sprinter. At launch, it won’t even manage half a percent of a gee. The Chinese ship took off twenty times faster, the rabbit to the American tortoise. It couldn’t keep that up. After a handful of hours of that, the Odyssey had exhausted its reaction mass and was coasting on its trajectory to Saturn—as it still is.
     “The Nixon is a marathoner. The nuclear-electric VASIMR system won’t shut down after a few hours or even a few days. It can run nonstop for months, accelerating the ship to the halfway point near Jupiter’s orbit and then continuously decelerating it until it arrives at Saturn. The VASIMRs will only add a handful of centimeters-per-second velocity to the Nixon every second. But there are a lot of seconds—more than eighty thousand in a day, two and a half million in a month. That adds up to a lot of velocity.”

---

     FLORA: Becca, you’re a nuclear power plant engineer, not a rocket scientist. Why are you part of this mission?
     BECCA: The propulsion system for the Nixon isn’t a conventional nuclear rocket like the Chinese are using—it is, in large part, an electric power plant. That’s where my primary responsibilities lie.
     FLORA: Can you explain the difference in the engines on the two ships?
     BECCA: Sure, the Chinese are using a nuclear thermal rocket (nuclear lightbulb). Take a nuclear fission reactor, get it real hot, and pump hydrogen through it. The hydrogen heats up and jets out the back, and there’s your rocket engine. The only difference between their engine and a chemical rocket is that they’re heating the gas in a fission reactor instead of by combustion.
     FLORA: Okay, and the U.S. mission is using . . . ?
     BECCA: An engine that’s never been used on a major space mission before, called a “va-si-meer” . . .
     FLORA: You better spell that.
     BECCA: V-A-S-I-M-R, which stands for “Variable Specific Impulse Magnetoplasma Rocket.”
     FLORA: Oh, well, that clears everything up!
     BECCA: [laughter] Well, then, my work here is done. Okay, before everyone tunes out . . .
     FLORA: It may be too late for that [more laughter] . . .
     BECCA: What we have is a really big, fancy kind of ion engine. We take a gas and ionize it—knock an electron or two off of each atom so instead of being electrically neutral, each atom has a positive charge. We funnel those ions into a big particle gun. It uses magnetic and electric fields to grab those ions and accelerate them to very high velocities. We squirt them out the back, and there’s your rocket exhaust.
     FLORA: So where does the reactor come in?
     BECCA: Reactors, actually. We have two. That way in case something goes wrong with one, the Nixon isn’t dead in space. The reactors come in because our rocket engine uses electricity. It takes electric power to ionize the gas and it takes electric power to generate the magnetic and electric fields that accelerate the ions. We use the reactors to power an electric plant, just as we would on Earth.
     FLORA: Then, the ship’s power plant is a lot like the systems you’re used to designing for electric utility companies?
     BECCA: Yes, which is why they want my expertise. But there are some important differences. The ship’s power plant is considerably bigger than any reactor complex in commercial use. It puts out nineteen gigawatts of thermal power. That’s an awful lot of power in a small space, and make no mistake, it is a small space. The reactors themselves wouldn’t fill up this room.
     FLORA: Wow. Why are reactors in regular power plants so much larger?
     BECCA: Partly because they don’t have to be any smaller. It’s a lot harder to manage that kind of power in a small volume than a large one. Also, the reactor on Earth is surrounded by tons and tons of shielding and containment vessels to protect people and the environment from its contents. In our spaceship, we dispense with most of that. The reactors are all the way at one end of the ship and just need a small shielding cap to provide a radiation shadow for the occupied portions of the ship. We don’t need a big, bulky containment vessel. If something goes wrong and the reactor fails catastrophically, that’s gonna be the end of us, anyway. My job, along with the reactor designers, is to make sure that won’t happen.
     FLORA: Okay, then, what makes a VASIMR better than a tried-and-true nuclear thermal rocket?
     BECCA: Two things—the first is that we can get a much higher exhaust velocit out of the VASIMR. That means we can go a lot faster using the same amount of reaction mass. I think I can explain how that works to your listeners.
     FLORA: I certainly hope so.
     BECCA: Imagine you’re on a pair of roller skates and you’re holding a bowling ball. If you throw the bowling ball away from you, you start rolling backwards. Action balances reaction, thank you, Newton. Now suppose instead of a bowling ball, you have a small pistol that you fire. What happens? The bullet goes forward and the recoil sends you rolling backwards. A very small mass, like that bullet, can push you just as hard as the bowling ball did, if it’s going very fast. The VASIMR gives us a lot more push than a nuclear thermal rocket would, for the same amount of reaction mass.
     FLORA: And the second thing?
     BECCA: This part’s a bit peculiar. The way rocket physics works, it takes a lot more energy to get the ship up to speed with a high-velocity exhaust than with low velocity. Our fast-moving ions are very efficient at using a small amount of reaction mass but very inefficient at using small amounts of energy.
     To put it another way, if we want to use as little reaction mass as we can, we want to throw bullets. But if we want to get the most speed from our power plant, we want bowling balls.
     In between those extremes is a happy medium. If we adjust the exhaust velocity as we go, we can get by with a ship that’s about half as big as it would be with a fixed exhaust velocity. Rocket scientists call the exhaust velocity “specific impulse,” which is where the first three words of the engine’s name come from: “variable specific impulse.”
     FLORA: On to my next question. If VASIMRs are better than thermal nuclear rockets, why didn’t the Chinese use them in their ship?
     BECCA: VASIMRs have one big disadvantage—they’re electric. A nuclear thermal rocket puts all its heat energy into the exhaust, it’s essentially one hundred percent efficient. We have to convert the heat of the reactor into electricity and using every trick we know, we can only do that with fifty-five to sixty percent efficiency. The other forty-odd percent? It’s waste heat, and if we don’t get rid of it the ship’ll fry. To give you an idea how much heat I have to deal with, it’s as if you took all the power used by a city the size of, say, Minneapolis, and stuffed it inside our little bitty rocket. My job is to get rid of it! Believe me, that’s not easy. There’s a lot of specialized plumbing, some humongous heat radiators, the whole thing is very complicated. Our early tests . . . well . . . we had some hiccups.
     FLORA: That radiator test in orbit last January?
     BECCA: Ummm, yeah, that didn’t go so well. But cutting-edge engineering is like that. That’s why we test instead of just flying off. Anyway, a gas core reactor with a hydrogen flow-through, like the Chinese use, is child’s play compared to this. That’s a lot more stable and we have a lot more experience with it. If we didn’t need to be going really, really fast, we’d never be using VASIMRs. The Chinese were originally going to Mars. They didn’t need to be going anywhere that fast. A simple nuclear thermal rocket would get them there in a few months with a very reasonable mass-to-payload ratio. To get that same ship to Saturn, though, they had to hot-rod the whole setup, add some truly monstrous hydrogen tanks for the additional reaction mass they need, and it’s still going to take them a year and a half to get there. They are the tortoise and we’re the hare. This time, though, the hare is going to win the race.

---

     A major flare would unleash a burst of X-rays, and at the Nixon’s distance (0.2 AU, inside orbit of Mercury), the hard radiation would hit them in a few minutes—most of the crew wouldn’t get enough warning to reach the safety of aft Engineering, where they would be shielded by the huge water tanks that provided reaction mass for the VASIMR engines.
     There were hidey-holes in each module of the ship, which, in a pinch, could accommodate the crew in a radiation-safe environment for the hour or so they might need protection—but it would be crowded and uncomfortable. Crowded and uncomfortable was better than dead.
     (Captain) Fang-Castro had insisted on drill after drill until every crew member showed they could make it to safety in less than ninety seconds, three times in a row. In the month leading up to perihelion, every crew member had come to hate the sound of the flare alarm.
     After the X-rays, there’d be a proton storm. The flood of charged particles was immensely damaging, biologically, but it would also wreak havoc with electronics, inducing massive eddy currents in anything metallic.
     The hidey-holes and the water tanks were enough to protect the crew but there was no way to electrically shield the entire ship. Shunts and circuit breakers would provide some protection, and the ship builders believed the craft would make it through without fatal damage. Nobody was quite sure, and there was no way to test for it.
     The worst possibility was that they’d be hit by a coronal mass ejection. If that happened, they were toast. The massive plasma stream would overwhelm any imaginable safeguards on the ship’s critical systems.
     There really wasn’t much to be done except prepare for what they reasonably could. Space weather could give them some advance warning, but the Nixon was not a maneuverable ship. The math was simple and irrefutable: the ship was barreling along at one hundred and fifteen kilometers per second deep in the sun’s gravitational well. At best, its engines could alter its velocity by two kilometers per second in a day’s time. Major course changes were out of the question. If the Nixon found itself on a collision course with a coronal mass ejection, then a collision was what was going to happen.

---

     Plaster-and-paint involved repair of collision damage, an ongoing issue for the Nixon. Most of the repair work was done by his own maintenance crews, but he trained extras in case there should be a substantial, but non-fatal, hit, which would require large crews both inside and outside.
     The Nixon was traveling a hundred times faster than a high-speed bullet. A millimeter grain of interplanetary sand packed the same wallop as a rifle slug when it hit the ship at a hundred and forty kilometers per second.

---

     The Nixon was ninety-three days out (about 3 months), eight hundred and ten million kilometers from Earth, six hundred and twenty million from Saturn, and on the far side of the asteroid belt. Becca rode a transport cart down the axle to the service egg bay, sucking on the morning’s third bulb of coffee as she went. The day before, the Nixon had reached the point where it would stop accelerating outward from the sun, turn tail-forward, and begin the three-month process of decelerating into Saturn.
     Becca had spent the day supervising the first controlled shutdown of the Nixon’s propulsion system since they’d left Earth’s orbit. Shutting down was less dangerous than start-up, but not a whole lot less tricky. Strictly speaking, it was unnecessary. The thrust of the four VASIMRs was low enough that firing broadside for a few hours would hardly affect their trajectory. But the techies running the simulations were antsy about those sail ribbons running at full velocity. The sims said they’d be fine in a rotating reference frame, but better to err on the side of caution and slow that molten metal down as much as they could.
     Engineering had to ramp down the output of the reactor, turbines, and generators while shutting down the VASIMR engines and slowly winding down the radiator system. Not too fast or the system would overheat; not too slowly or the heat exchangers would freeze up.
     The shutdown wasn’t complete. Becca didn’t want to restart the radiator system from scratch; cold starts were always rough. Happily, there was no need to. The reactors could churn out power indefinitely, and she’d had them wound down to the point where they’d be supporting the radiators at minimum output. She could bypass the main turbines and generators entirely, run the auxiliary power system—more than enough to maintain ship’s power and the heater and control systems for the heat exchangers and radiators—and toss in a bit extra to give the molten metal ribbons something to do as they cycled from nozzle to collection boom and back down into the heat exchangers.

     The process wasn’t all that difficult and they’d practiced it in Earth orbit. This wasn’t for practice, though, and there were no rescue ships if something went wrong.
     It all went smoothly, though, and the Nixon went into free fall.

     At that point, the attitude thrusters went to work, a complex and delicate orchestration of impulses that slowly rotated the entire structure—booms, struts, axles, and modules—a hundred and eighty degrees, so the engines were pointing away from the sun, and what had been the forward part of the ship was now facing back the way they came.
     The crew wasn’t overly worried about impacts on the engines. Their cross-section was small, and they were now well clear of the asteroid belt; in fact they were approaching Jupiter’s orbit. Jupiter, fortunately, was far away. Becca didn’t have to be concerned with the Jovian gravity well or the massive radiation belts. Even Jupiter’s leading and trailing Trojan asteroids were far off: their orbit simply made for a mental benchmark.
     The next one would be the rendezvous with Saturn.
     On this day, she and her crew would bring the engines back online, essentially, shutdown in reverse. Again, as they’d practiced so many times in Earth orbit, it would be a slow and coordinated ramping up of reactors, turbines, generators, engines, while bringing the radiator system up to full speed.

     If everything went as planned, they’d be at Saturn in a little over three months.

---

     The Nixon, with its near-kilometer-long radiators and three-hundred-meter main axle, was larger than the Odyssey, but it was like a box kite made of balsa wood and string, long thin columns and beams tied together with graphene guy wire. The Chinese ship was only two-thirds the size of the Nixon (perhaps 225 m long?), but it looked like a tank (because it has twenty times the acceleration and needs twenty times the structural strength).

      In our book, the Chinese are using tried-and-true technology, at least for fifty years from now. Nuclear thermal rockets can get much higher exhaust velocities than chemical rockets (see NERVA, Wikipedia). The Celestial Odyssey uses an exotic reactor design called a “lightbulb reactor” that no one’s built yet but that engineers have designs for. If we needed them, we could have them in fifteen or twenty years. The Chinese ship pushes that technology to the limits of what engineers think we can do. It heats the exhaust to around 9000°C and it uses hydrogen for the reaction mass. It could use any gas as the reaction mass, because it’s just heating it up in a reactor, not trying to burn it. Hydrogen is best because lighter atoms move faster than heavier ones at the same temperature, and hydrogen is as light as we can get. That gets the Chinese the highest exhaust velocity, 22 km/s, which is five times better than burning hydrogen and oxygen.
     With that kind of exhaust velocity and a mass ratio of 7 or so, the Chinese can get from Earth to Saturn in around a year and a half (delta-V about 42,800 m/s). (How do we know that? Later.) It doesn’t get them back, but Saturn’s rings are water ice and the Chinese can break that down to get the hydrogen to refill their tanks.

     Could such a spacecraft make it to Saturn in half a year? Not likely. They’d need nearly three times the delta-vee, and the mass ratios would be hundreds to one. The Americans need something better. Enter the VASIMR engines. VASIMR stands for “Variable Specific Impulse Magnetoplasma Rocket.” “Specific impulse” is how rocket scientists refer to exhaust velocity. We didn’t make the VASIMR up. They’re being tested on Earth, fairly small ones. Ours are a lot bigger, and a little better-performing, but it’s fifty years from now. Building bigger and better VASIMRs doesn’t look hard; powering them does, and we’ll get back to that.

     Becca’s Science Friday interview in Chapter 21 explains why variable exhaust velocity is, in general, a good idea. The Nixon can get more thrust for the same amount of power, when the ship is fully laden, by keeping the exhaust velocity low. That consumes more reaction mass for the same delta-vee, but it gets the ship up to speed faster, shortening the trip time. As the ship gets lighter, it can get by with less thrust and still keep up the acceleration, so the Americans can run the exhaust velocity higher and make more efficient use of the remaining reaction mass.
     We built a spreadsheet to let us play around with different velocity profiles. For a trip time of four to five months, we were able to get the ship down to a mass ratio of 10 with an exhaust velocity that varied from 35 up to 300 km/s. That’s about half the mass ratio we could come up with for a fixed-specific-impulse ship of any remotely plausible design. Go, VASIMRs!

     The reason the Americans’ ship uses water instead of straight hydrogen is because it doesn’t need to use straight hydrogen. The exhaust velocity that comes out of a VASIMR depends on the charge on the ion and its mass. Hydrogen produces the highest exhaust velocity. Strip off one electron and you’re left with a charge of one and an atomic weight of one. Strip one electron off of oxygen and you’ve got a charge of one but an atomic weight of 16, so the electromagnetic fields in the VASIMR won’t push it anywhere as fast. One of the ways the Nixon has to tailor its exhaust velocity is to tailor the mix of oxygen to hydrogen.

     So how can the VASIMR keep up with a “lightbulb” ship? The lightbulb gets its initial velocity from one long burn—the rest of the trip it’s in free fall, until the very end, when another burn will slow it down so it can go into orbit around Saturn. With a VASIMR, you simply don’t turn it off. You’re making a much more economical use of your reaction mass, and the accumulating thrust eventually adds up to much more velocity than is possible with a lightbulb.

     VASIMRs have a problem, though. They’re powered by electricity, lots of it. The only way we know of to generate so much power is a nuclear power plant.
     Surprisingly, the reactor isn’t the problem. Reactor cores can generate amazing amounts of thermal power. You have to get that heat out or the core melts down, but NASA figured out how to build a liquid-lithium-cooled core the size of a coffee can that would output 2.5 megawatts back in the 1970s. That’s already as good, in terms of both watts per kilogram and watts per cubic centimeter, as what the Nixon needs. The two reactors in the Nixon are each four thousand times bigger . . . but they are not better.

     The huge problem the Nixon faces is that only a little more than half of that heat can get converted into electricity that goes into thrust and is kicked out the back in the VASIMR exhaust. There are some fundamental thermodynamic principles that make it unlikely we’ll ever be able to do much better than that. The rest of the heat, nine gigawatts or so, ends up being waste heat and has to be disposed of before everything melts down. That is the really, really hard problem in space.
     There are only four ways to move heat around: convection, conduction, transport, or radiation. The first two can’t get the heat off of the ship—in a vacuum, there’s nothing to conduct heat away from the ship, nor is there gas circulation to convect it. You could transport it off, use it to heat up reaction mass and send it out a rocket nozzle. But that’s just a nuclear thermal rocket like the Chinese have, and it’s nowhere near efficient enough for the Nixon. It also requires infeasibly humongous amounts of heat-absorbing reaction mass; nine gigawatts is an awful lot of heat to be dumping off the ship, day in and day out for months and months.
     We’re left with radiation. Simple physics describes getting rid of heat by radiation (see “blackbody radiation,” Wikipedia). At 80°F (27°C or 300 Kelvins) a two-sided radiator can dump about a kilowatt of heat (1 kW) per square meter into space. That’s a lot by human standards, but it means the Nixon would need about nine square kilometers of radiator to get rid of all its waste heat—roughly the area of 1,700 American football fields. That would weigh far too much. Physics works in our favor, though. The amount of power you can radiate goes up as the fourth power of the Kelvin temperature. At 600°C (870K) that same square meter of radiator can dispose of 65 kW of heat. That takes the radiator down to a manageable size.
     This requires running the whole power plant hotter, because the theoretical efficiency of the power plant is determined by the starting temperature vs. the final temperature (see “Carnot,” Wikipedia). We’ve kicked the final temperature up nearly threefold, so the source temperature has to go up accordingly. That takes us out of the range of boiling water reactors and generators and into the world of pressurized liquid sodium, running at a red heat.
     Is this insane? Definitely, by today’s standards. We looked at power plant performance and benchmarks for the past century-plus, and extrapolated extrapolated those trend lines forward fifty years. In half a century it doesn’t seem so crazy.

     What kind of radiator can run at 600°C? The only possible working fluid is molten metal, and that also works in our favor. Melting metals can absorb huge amounts of energy, far more than any other material. Aluminum, for example, melts a little higher than our working temperature, and melting 2.5 kilograms of it per second will suck up a megawatt of heat. Our imagined radiator alloy is a little better than pure aluminum, but not a whole lot. It’s all plausible; we’re still doing science and engineering, not wild imagining.
     Some simple algebra and punching numbers on a four-function calculator tells us the size for the radiator sail, the speed of the sail ribbons, and the thickness of the ribbons. They are interrelated. The speed of the ribbons has to allow enough time for the semi-molten ribbons to completely solidify before they reach the collection beams. The amount of ribbon being extruded each second has to be sufficient to carry all the waste heat from the Nixon.

     A minor bit of cleverness. That thin layer of solid cladding that’s frozen onto the ribbon as it’s extruded does three things for us. It roughens the surface, which improves the radiation characteristics (see emissivity). It keeps the ribbon a ribbon—molten metal has very fierce surface tension, and a thin liquid ribbon would otherwise break up into droplets almost immediately. Finally, it prevents evaporation of the metal. Typical metal vapor pressures are very low at 600°C, but you’d be talking about over a quarter kilometer of surface area and a year’s time. You’d have to carry along considerable amounts of spare radiator metal to make up for evaporative losses. In space, it’s waste not, want not.

     So this is all pretty great: we have two ships that will fly more or less the way we need them to. What will those flights look like? That’s orbital mechanics, and while the physics of that is high school–level Newton’s laws stuff, the calculations are lengthy. They would’ve been impossible without personal computers.
     We got lucky on one point. The starting date for the book was an arbitrary choice and the pacing up to the point where the ships launch was dictated by what would keep the story moving properly (and launch windows—the right alignment of the earth, sun, and Saturn). As it happened, we wound up in the easiest time frame for calculating orbits. None of the other planets came anywhere close to possible orbits for our spacecraft. All we had to deal with was the earth, sun, and Saturn.
     We did the orbit simulations in a Windows program called Orbiter. That wound up consuming hundreds of hours of computer and personal time. Continuous-boost spacecraft like the Nixon are still exotic enough concepts that very few software packages can deal with them easily. Those programs are expensive and specialized, their normal clientele being major aerospace companies and research institutions.
     Orbiter could very accurately calculate the motion of the spacecraft for each step along a trajectory, but those steps had to be carefully and manually specified. When the Celestial Odyssey was falling free, far from the gravitational influence of the earth or Saturn, a step could span months of simulated time (on the computer, that ran in minutes). We just needed the steps to be close enough together that we’d know where the Chinese were at any particular time in case that needed to be mentioned in the story. Entering or leaving orbit involved a lot more steps because the ship’s trajectory was changing very quickly.

     Simulating the Nixon was much harder. Except when Bad Things happened to it, it was never in free fall. The simulation steps had to be close enough together to accurately model the trajectory of a spacecraft that was always accelerating or decelerating. Spiraling out from Earth orbit took about fifty steps to simulate, similarly for the transit from Earth to Saturn. At each step, the thrust parameters were keyed in, the time interval specified, the step calculated, and the data—velocities and positions relative to the earth, sun, and Saturn—extracted and transferred to a spreadsheet. That spreadsheet also produced a plot for us showing the spacecraft’s trajectory against the earth, sun, and Saturn. A different spreadsheet tracked energy budget—reaction-mass consumption, mass ratios, exhaust velocities, and accelerations.
     We ran nearly a hundred such simulations as we developed the story. Some of those runs were dictated by plot changes as we developed the novel, but most were experiments designed to answer questions. How does the Nixon’s departure date affect the travel time? How long can the Chinese wait before chasing after the Nixon? What’s the best way to get back to Saturn? What’s the most efficient trajectory for a return to Earth?

     Real rocket science. We think we did it pretty well.

Another ship that will give old rocket fans a sense of haunting familiarity. Whether you saw it in the old Life Science Library volume Man in Space or as the Project SWORD toy, it is another bit of your childhood that would actually work.