(ed note: Nuclear Salt Water powered Frigate New Jersey wants to dock with colony on asteroid 624 Hektor. Nuclear Lightbulb powered Tug Brutus assists. Remember, with nuclear drive rockets using shadow shields the only safe way to dock is nose-to-nose.)

"Conn, tug Brutus requests permission to dock."

"Permission granted. Close the fuel valves."

As New Jersey shut down her main engine, the powder blue UN tug approached on thrusters. Fitzthomas watched on external cameras, never liking it when his ship was in others' hands, but powerless to do anything about it. The traffic rules within 100 k-klicks of Hektor were as plain and draconian as those of LEO: no unauthorized burns, and no open cycle nuclear reactors in operation under any circumstances. And the UN tug drivers were pretty good.

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"Conn, Brutus is making her final approach. Docking in one minute."

"XO (executive officer), signal all hands."

"Aye sir." A moment later, Allen's voice came over the shipwide intercom. "All hands, prepare for ship to ship dock." He switched back to his private channel to the captain. "Let's hope whatever blue hat is driving that donkey cart knows how to actually fly a spaceship."

As it turned out, he could. Brutus and New Jersey touched nose to nose at barely a meter per second. Heavy duty docking latches locked the ships together, and Brutus extended a fiber optic probe into a receptacle inside New Jersey's docking ring.

"Conn, we have hard dock with Brutus. They are requesting propulsion access."

Fitzthomas removed a hard blue plastic key from his extensive key ring and inserted it into a blue lock. "Transferring propulsion access on my mark....mark." He turned the key and the clear plastic ring around the lock lit up blue. Brutus's pilot now had control over New Jersey's maneuvering thrusters. For the duration of the ride, Fitzthomas's hand would hover over the key, ready to cut Brutus out of the system, just in case.

Brutus's pilot could now also see what New Jersey's docking radar and aft cameras saw, a necessity since the bulk of the frigate would block the tug's line of sight—visual or otherwise—while they maneuvered. Brutus was little more than a nuclear engine (closed cycle, which meant no reactant could escape with the exhaust, at the cost of about half the ship's specific impulse, which didn't matter for a short range, low speed tugboat) with a sturdy docking ring and a cockpit jammed in between. In some operations, the tug was unmanned and guided by computer or radio control operator until it docked with the target ship, and then the tug would relinquish control of its systems to the target ship's pilot, or a harbor pilot brought on board for the purpose, saving some of the mass of reactor shielding for the cockpit and some of the ulcers for shipmasters, but the UN insisted on doing things the old fashioned way. Fitzthomas had heard of tugs which did the entire trip by remote control or onboard computer, but he'd rot in hell before he let some toe-picker play model airplane with his spaceship, and the United States Space Guard agreed.

New Jersey's thrusters fired, heeling the ship around 180 degrees, so that her bow, with the tug latched on, faced Hektor. Brutus then fired her main engine, continuing the deceleration burn New Jersey had started. Fitzthomas felt himself lift ever so slightly out of his chair. His ship's decks were aligned so "down" when they were under thrust was towards her main engine. With Brutus docked with her nose and her thruster pointing the other way, "down" was now, temporarily, in the direction of the overhead. One more thing for Fitzthomas not to like about the tugboat.

"Approaching Hammarskjöld docks. Ten minutes to docking."

1.2.3.2 Nuclear-Electric Tug Concept

     During the July 1976 NASA/Boeing working session, discussions were held with Mr. John Steams of JPL on the subject of nuclearelectric tugs (NET'S). Sizing and performance estimating data were obtained as reflected in the analyses described below.

Argon MPD Thruster-Experiments at Princeton have lased argon magneto plasmadynamic (MPD) devices to create plasmas for plasma physics experiments. Analyses of similar devices used as thrusters indicates that dearable performance characteristics may be obtainable. Table 1.2-1 indicates representative target values. These performance figures have not been confirmed by test, but we judged to be reasonable w3en associated with a system like the NET requiring a major development effort.

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Reactor Power Generation Systems-The reactor design approach, suggested by Mr. Steams, employs a high temperature, fast spectrum reactor cooled by heat pipes. Either thermionic or high temperature Brayton conversion systems are potentially practical.

Reactor core assumptions are given in Table 1.2-2.

Power and efficiency budgets are given in Table 1.2-3 and 1.2-4 for the thermionic and Brayton systems.

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The thermal power and efficiency assumptions were used to derive the reactor parametrics show in Figures 1.2-7 and 1.2-8.

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Figure 1.2-9 illustrates the fuel form concept for the thermionic system including heat pipe. Dimensions shown are representative. The thermionic system requires only a single active loop, for heat rejection.

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The cycle concept is shown in Figure 1.2-10. Also shown is the Brayton cycle schematic with a primary liquid rnetal loop coupled to the typical closed Brayton cycle.

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Radiator area requirements are shown in Figure 1.2-11. It is anticipated that the Brayton cycle system will require emergency heat removal radiator in the primary loop. The emergency heat removal is required only to handle after-heat and the radiator could be quite small.

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Gamma shield dimensions are given in Figure 1.2-12. For the Brayton system the gamma shield encloses the entire primary loop to minimize radiation from neutron activation of the primary loop fluid.

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Figure 1.2-13 shows outer gamma shield specific mass assuming 100 g/cm² (205 lb/ft²) shielding. The gamma shielding is intended to allow unimpeded manned operation around the NET with the reactor shut down. The shielding allowance is a rough estimate; the manned operation assumption was arbitrary but appears reasonable for the intended mission use. Very little neutron shielding is provided; the unshielded standoff distance for manned operations with the reactor at full power will be on the order of 10 km (5.4 n mi).

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Specific mass estimates for other items are given in Table 1.2-5 and 1.2-6. The Brayton power processing mass is greater than that for thermionics because AC/DC conversion is required. The thermionics radiator is estimated heavier per unit area because it includes a complete licluid loop; the Brayton gas radiator includes gas inventory and containers.

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A mass growth allowance of 25% was applied for connecting structure, auxiliary propulsion, avionics, controls, and unidentified items. The completed specific mass estimates, excluding main propellant tankage, are shown in Figure 1.2-14.

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Performance and Design Point Selection-These specific mass estimates were used to develop the performance estimates shown in Figures 1.2-15 and 1.2-16. Mass of main propellant tankage was estimated as 5% of the main propellant (argon) required. Since all the power generation systems were similar in specific mass, average values were used for the performance calculations

The geosynchronous space station delivery mission requires delivery of 6l,000 kg (135,000 lb). The 2 megawatt (jet power) tug can perform this delivery in about 75 days, or can deliver the entire station as a single payload in about 110 days

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Reactor disposal can be accomplished by sending the NET to solar system escape (no payload). Propellant required for the 2 MW_j NET is 120,000 kg (265,000 lb): this requirement sizes the propellant tank. 2 MW_j with a single-ended reactor and high temperature Brayton cycle was arbitrarily selected for a point design. The double-ended thermionic system is essentially equivalent in performance: a tradeoff beyond the scope of the FSTSA study would he required to make a selection. Design data are summarized in Table 1.2-7.

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The point design is shown in Figures 1.2-17 and 1.2-18. A mass estimate for this point design is provided in Table 1.2-8.

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Take the first trip: Tug with just the core (no auxiliary units) for 1,800 kg dry mass, 14,700 kg of propellant, and 2,000 kg payload. First leg of the trip is traveling from LEO to GEO, requiring 4,400 m/s ΔV.

R = e^{(ΔV / V_e)}

where

ΔV = ship's total deltaV capability (m/s)
V_e = exhaust velocity of propulsion system (m/s)
M = mass of rocket with full propellant tanks (kg)
M_e = mass of rocket with empty propellant tanks (kg)
R = ship's mass ratio
e^x = antilog base e or inverse of natural logarithm of x, the "e^x" key on your calculator

Plugging in our values:

R = e^{(ΔV / V_e)}
R = e^{(4,400 / 4,410)}
R = e^(0.9977)
R = 2.7

We can use this mass ratio to calculate how much propellant was burnt. Mass ratio is

R = M / M_e

Simple algebra gives us:

M_e = M / R

Plugging in our values:

M_e = M / R
M_e = (1,800 {dry mass} + 14,700 {propellant} + 2,000{payload}) / 2.7
M_e = (18,500) / 2.7
M_e = 6,850 kg

So the tug arrives at GEO with a total mass of 6,850 kg. We know that 1,800 kg is dry mass and 2,000 kg is payload. By subracting we see that there is 3,050 kg of propellant left.

In GEO, the 2,000 kg payload is delivered, and is no long part of the tug's total mass.

It is time for the second leg of the trip. Now the tug has a wet mass of 1,800 + 3,050 = 4,850 kg. It has a mass ratio of 4,850 / 1,800 = 2.69. How much ΔV does it have?

ΔV = V_e * ln[R]
ΔV = 4,410 * ln[2.69]
ΔV = 4,410 * 0.99
ΔV = 4,370 m/s

4,370 m/s is close enough for government work to the 4,400 ΔV required to travel from GEO back to the LEO parking orbit.

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Gas-Core NTR COTV
Engine	Nuclear Lightbulb
Isp	2,250 sec
Exhaust Velocity	22,100 m/s
Thrust	890,000 N
Mass Schedule
Stage Element	Weight kg
Structures and Mechanics	18,780
Main Propulsion	56,850
Auxiliary Propulsion	600
Avionics	260
Electric Power	480
Thermal Control	1,220
Growth (15%)	11,730
DRY WEIGHT	89,920
Other Propellants and Fluids	2,000
INERT WEIGHT	91,920
Mainstage Propellants LOX LH2	124,280
STAGE WEIGHT	206,204 216,200

The nuclear Cargo Orbital Transport Vehicle (COTV) concept analyzed combined the desirable features of the chemical COTV and the electrical COTV — high thrust and high specific impulse, respectively. The stage, shown on Figure A-21, has a nuclear gas core, light bulb-shaped engine with a theoretical specific impulse of 2250 seconds and a thrust level of 890,000 newtons. The component mass breakdown is given in Table A-3.

Although such a system could meet the short trip time requirement for personnel transfer and the high performance requirement for cargo transfer, the development risks and the presence of nuclear materials in LEO eliminated this system from further consideration.

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Gas-Core NTR COTV
Energy Source	Uranium 233 dioxide
Propellants	LH2
Specific Impulse	2080 to 2425 s
Thrust	445 to 1780 kN
Engine Mass	42,000 to 91,000 kg
Pressures	271 kN/m2 (operating) 384 kN/m2 (maximum)
Mass Schedule
Stage Element	Mass kg
Structures and Mechanisms	18,780
Main Propulsion System	56,800 56,850
Auxiliary Propulsion	600
Avionics	260
Electric Power	480
Thermal Control	1,220
Growth Allowance (15%)	11,730
DRY MASS	89,920
Auxiliary Propellants and Fluids	2 000
INERT MASS	91,920
Mainstage Propellants LM2 LH2	124,290
OTV TOTAL MASS	216,210

3. Gas-Core Reactor OTV

     A nuclear-reactor heat source was considered as an alternative to the solar array to power the OTVs (orbital transfer vehicles). The gas-core reactor was studied as the concept most adaptable to this mission and is presented in Figure 33.

     The specific impulse of a nuclear propulsion system is intermediate between that of chemical systems and electrical propulsion systems as indicated below:

• LO₂/LH₂ ∼470 sec
• Nuclear rocket ∼2,000 sec
• Electric rocket propulsion ∼6 to 8,000 sec

     Mass in orbit, hence cost, can be expected to be less with higher Isp.

     Neutron and X-ray radiation shielding is required for reactor usage in proximity to personnel. This consideration would seriously limit the flexibility with which such a vehicle could be used. Shielding is heavy and shielding design is a difficult problem. After-heat disposal during reactor shut-down is also an important consideration. Unshielded reactors, on the other hand, would require remote handling so that malfunction repair and maintenance in space could be expected to be very difficult; however, it warrants further consideration.

     The basic concept of the gas-core reactor relies on the use of thermal radiant energy transfer from a high temperature (∼80,000° K) radiating fissioning uranium plasma to a submicron tungsten particle-seeded hydrogen propellant stream. The plasma is vortex-confined by a cool nonabsorbing buffer gas. In one of the several gas-core reactor concepts which have been conceived, the fuel and buffer gas flows are separated from the propellant stream in the core by a transparent wall which allows containment of the fuel within a closed-loop circuit. The fuel would be processed for subsequent reinjection into the core region. Propellant exit temperatures in the range of 4,000° K to 6,700° K are predicted for the previous range of fuel-radiating temperatures. Corresponding specific impulse in the range of 1,000 sec to 1,900 sec and thrust-to-weight ratios of 0.3 to 1.3 have been estimated for engine powers of 600 mW to 4,600 mW. (Engine mass without propellant is 39,000 kg or 85,800 lbm.)

     The gas-core reactor engine offers the combination of high thrust and moderate specific impulse with the result that rapid LEO-to-GEO trips can be made. Thus, perhaps as few as one vehicle would be required, consequently reducing mass in LEO. However, it must be realized that crew shielding (shallow (sic) shielding) must be incorporated that, depending on the safety considerations, will add to the engine basic weight. An assessment must also be made of potential upper atmospheric pollution.

     The technology development for the gas-core reactor would probably be longer than electric propulsion devices, but the high thrust, high specific impulse combination may make the gas-core reactor a promising candidate for use in applications beyond the initial deployments.

     Although the gas-core reactor requires advanced development of several disciplines, numerous "proof-of-principles" experiments have been conducted over the past 15 years. For instance, a seeded flowing gas stream (simulating the propellant) has been heated by radiation from a dense plasma to temperatures exceeding 4,000° K. A radiating plasma (equivalent black-body temperature of 6,000° K) consisting of argon and UF₆ (uranium hexafluoride) has been successfully contained within a container of cooled fused silica without causing coating of the walls and transmitting over 90 per cent of the source radiation through the walls. A system was developed that permitted separation of the uranium from the argon and demonstrated that recirculation of the UF₆ was indeed feasible.

     While these experiments have been on a small scale relative to that required for the nuclear light bulb engine, they do demonstrate that much of the technological "know-how" necessary has been developed.

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     So four days ago or there abouts, I put a poll up on Google+ with a selection of spacecraft I was thinking about making isometric cutaways of. The frontrunner is the Intra-Fleet Space Tug. That means, RocketFans, that we’ve got ourselves a project!

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     The context for this particular spacecraft, like the Cygnus capsule I also put in the poll, is the care and feeding of the distributed-network fortification that is a deployed UN Constellation in the Conjunctionsetting. In summary, the fleet’s configuration is a tetrahedron in space with a single control ship at the apex, patrol craft making up the other three vertices, and edges three hundred thousand kilometers long. Just how do you supply ships that are as far out as the Moon is from LEO?

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     In the article about how fleets work, I stated that the crews on the patrol craft could be swapped out by ferrying fresh people out via the Cygnus. While this would certainly work for crew transfers, you’d also have to detail additional craft for cargo transfers, of consumables and (if armed with rail guns) ammunition. As versatile as the Cygnus is, it cannot not re-supply that most important consumable resource in terms of tactical movement, propellant.
     To put the problem into perspective, a Cygnus stack is a rough cylinder 4.5 meters in diameter and about ten meters long. The propellant tanks on a Type A Patrol Cutter are 8 meters in diameter and total thirty meters long. And there are two stacks. Clearly, to refuel a patrol ship, we need a real tanker.
     I’ve said it before RocketFans, and I’ll surely say it again: Atomic Rockets is an invaluable resource for the budding rocketeer. The “Realistic Designs” sections are a veritable clearinghouse of old NASA designs that were pretty good but never got a decent budget. These oldies make for a great library of inspiration when designing any spacecraft that is meant to work with real-world physics. For our Intra-Fleet Tug, I was inspired by the Johnson Space Center’sTug study, who’s image I used in the Poll. This beauty is a two-stage ferry to get from LEO to GEO where NASA was going to build a solar power station.
     Anyway, a light-second is good deal further than the LEO/ GEO distance, right? In kilometers, yes, but in Delta-V, not even close. It takes a whopping 4.33 km/s to go from LEO to GEO, but a paltry 2.74 km/s to get from LEO to Lunar orbit…a little over a light-second away.
     Gravity is funny like that.
     So our tug only needs about 75% the range of the JSC version. Since that design was staged and the first staged carried the spacecraft 85% of the way to GEO we could just lop of Stage I and call it a day. But where’s the fun in that?
     The problem with just ripping of the JSC design is that it isn’t a tanker. We need to be able to deliver a large amount of propellant, so we’re going to need a large spacecraft. Something that could haul at least a quarter or half of the Delta-V needed to completely refuel a Patrol craft. What follows is an experiment: I’m thinking of just taking an entire rocket stack from a Patrol craft and slapping a command module on the front for our Tug. Let’s see how that would work, shall we?
     First of all, we need to dust off our rocketry equations so we know what variables we need to consider. We’re going to need to know the Tugs dry mass, wet mass, and engine details such as propellant flow, thrust, and exhaust velocity. Since we’re using the dimensions of the propellant tanks from the Class A Patrol Craft, and possibly one of its main engines, that gives us a great place to start. In fact, lets crunch the numbers for the Patrol rocket’s main engine and an alternate, say something along the lines of the J-2 from the Saturn V’s SIV-B stage.
     First, let’s establish the tonnage for the Tug without it’s engines. We’ll want a decent sized crew module, because gaming, and also so we can have cadets aboard during all flights. In Conjunction, like in Heinlein’s Space Cadet every UN convoy and spacecraft has a group of peacekeeper candidates learning how to work in space by working in space. I see an actual crew of about four: a Flight Commander (F-Com), Guidance Procedures Officer (GPO), Maintenance, Mechanical Arms, and Crew Systems Officer (MACS), and a Payload Officer (Payload). Add as many again of Candy-Cruisers, and you’ve got eight people in the command module. That’s a bit crowded for a Tug, but we can use hot-bunking with to limit the sleeping berths to four. The CM must also have at least a pair of robotic arms, and a sturdy docking module for carrying passenger capsules and cargo pods. Behind the CM will sit a flared-out service module, with avionics, life support, and computer systems. The SM will be mated to a 30 × 10 meter saddle truss, which is what will actually hold our propellant tanks and provide a mount for the rocket stack. But in addition to all of that, we will also need a passenger module and cargo pods, so we need to know the mass for all of those as well.
     Here’s how it breaks down:

System	Mass (kg)
Command Module	12,671
Service Module	3,000
Saddle Truss	24,119
Propellant tanks	24,119
Passenger Module	7,540
Crew Average Mass	2,400
Cargo/consumables	392,883
TOTAL DRY MASS	466,732
LH2	71,204
LOX	305,788
TOTAL PROPELLANT MASS	376,992
TOTAL WET MASS	843,724

     I arrived at some of these number dubiously, so take them with a grain of salt. The CM mass is from the Trans Hab Calculator on the AR website, the SM is from the JSC Tug, the truss is simply repeating the mass of the propellant tanks, since I couldn’t find any reliable numbers for that. The Passenger module is also from the JSC tug, while the consumables and cargo masses are calculated for the tugs trip out and back, as well as 30 days of supplies for the 20-person crew of a Patrol craft. And of course, we can’t forget the mass of the crew and passengers themselves, plus what ever possessions they can carry inside their regulation 100 kg mass-limit. Finally, the propellant tank mass is 6% of the propellant mass, as per Dr. Rob Zubrin, and the propellant masses came from the Useful Tables appendix from Atomic Rockets. But the most important thing to remember is that we have no engine yet.
     The Class A Patrol craft uses an easy to maintain in freefall analog of the Space Shuttle Main Engine (SSME) so I could simply steal copy the vital statistics. Engine List on Atomic Rockets has these available. Just below that entry is the stats for the Tug engine we will also use. These are not exactly the J-2 stats, but they are for a NASA tug, and they have the information I need to calculate with, whereas sources on the J-2 did not.
     What we want to know is, assuming a 100-hour flight time, is how much propellant will be left in the big tanks at the end? We need to have spend no more than 1/3 of our propellant mass in transit. That way, we can refuel with another third (plus a bit extra) and use the remaining less-than-a-third to take our much less massive tug home.
     This means math. So, so much math.
     Well, not so much, perhaps. We know all the vital statistics for our engines, our mass numbers, our Delta-V budget, and our distances. By establishing an arbitrary travel time of 100 hours, we also provided a much-needed value for equations, and more important, the mass of needed consumables.
     An Intra-Fleet Tug that uses a “F-2b” SSME-analog will have a wet mass of 846,901 kg, or 847 tons. Let’s see if we can get from point A to B while only burning through 125,664 kg of propellant.
     Simple, right?
     If only using 125.6 tons of our propellant, we will be operating with a mass ratio of only 1.8 By using the Delta-V equation of Delta-V = Exhaust Velocity × ln(Mass Ratio). This results in a Delta-V of 2621.96 m/s, or 2.62 km/s. We need 2.74 km/s to get to our destination, so it’s close, but no cigar.
     If we attempt the same thing with our J-2 analog, we have a wet mass of 845,512 kg. This gives us a mass ratio of 1.8 again. However, the exhaust velocity is 4159.4 (I had to calculate it using the specific impulse, but that’s why we have algerbra in the first place). With the mass ratio and a lower exhaust velocity, the Delta-V is 2.45 km/s. Both engines are pretty comparable, but neither will get us out a light second and back.
     Or will they?
     The moon averages 384,000 kilometers from Earth. A light-second is only 300,000 kilometers. We actually have less distance to travel, and hopefully less Delta-V, than the 2.74 km/s we’ve been using. Possibly a lot less.
     I forgot that moving around a fleet formation like this is not remotely the same as moving around orbits. Moving from LEO to Luna is a Hohmann trajectory, which is a change between orbits from around one body moving at one speed to another body moving at a very different speed. When deployed, our constellation is all moving at a constant speed along a constant orbit/vector. This means that all spacecraft in the formation are at rest relative to one another. So we need to go from a starting velocity of (relatively) zero to a certain speed, coast, flip, and then decelerate back to zero. This is just a simple physics problem.
     This is also where our arbitrary 100-hour travel time comes in. With time and distance known, as well as acceleration (Thanks to the engine stats) we can solve for velocity and begin to figure out what we need to know.
     Solving the displacement equation gives us an average velocity of 833.333 m/s to travel a light-second in four days and change. This means we need a final velocity of 1666.666 m/s. Our SSME engine will take only 721 seconds to boost our monster tug to speed, and the same to decelerate at the other end. Now for the biggie – mileage. By which I mean, just how much propellant did we use up in those 1442 seconds?
     Turns out that’s an easy one, because we know the mass flow. A single SSME tosses 409 kilos out the back every second, so our Tug will have to burn 589,778 kg. This is more than the entire wet mass of the tug, so say nothing of the “one-third” we wanted to get by with.
     As for the J-2, we need to re-do our acceleration calculation so we can figure our burn duration. Unfortunately, with a burn duration of 1282 seconds one way, the performance is even worse.
     What went wrong? This tug has half the power or a patrol rocket – it should have at least comparable performance.

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     Having gone back over my notes I discovered my problem, and it’s an embarrassing one. The Class A Patrol Craft I just mentioned, the one that’s over twice as large as this tug? It has a dead weight tonnage of 70 tons. That’s it. The Tug has a dry mass of 466 tons. Well, there’s our problem!
     I designed the Patrol Craft to take into account the likely progression of materials science toward ever lighter and stronger materials. It was built out something that has the same strength of titanium, and half the mass. Add to that it’s outer skin is mostly carbon and aerogel – literally the least dense substance there is – and its easy to see that simply cribbing numbers from a design made when aluminum was the lightest thing you could build spacecraft of is a problem.
     Let’s try this again shall we?

System	Mass (kg)
Total Structure Mass	24,119
Crew Average Mass	2,400
Cargo/consumables	4,245
TOTAL DRY MASS	30,764
LH2	71,204
LOX	305,788
TOTAL PROPELLANT MASS	376,992
With J-2	409,544
TOTAL WET MASS	407,756
With SSME	409,337

     I not only went back and recalculated the structure mass using 22^nd century materials, I also hand-calculated the mass of the consumables and cargo, using NASA rations. Much better results. With these stats, the Tug can pull 4.43 m/s, and only has to burn for a total of 376, instead of 1442. This means we only burn 141,514 kg of propellant. With less thrust and more mass, I don’t feel a need to calculate for the J-2. 141.5 tons of propellant is 37% of our propellant mass. For the return trip, we’ll need less propellant, say, 25%? The Tug would only mass 126 at that propellant fraction, and accelerate at a whopping 14.4 m/s, or 1.4 gs. It will only have to accelerate for 115 seconds and burn only 43 tons of propellant, while carrying 96 tons. This is over a 100% reserve, enough that we could add another 20 tons or so to the 124 tons our Tug is pumping into the Patrol craft.
     So, there you have it, RocketFans, a glimpse into the hair-tearing-out, thankless job of designing a realistic spacecraft. I’m glad I just have to make these look good on paper. But the important part is, I can now draw a spacecraft with all the particulars I wanted to, and it will not only look realistic, it will be realistic. It’s capabilities and limitation will suggest numerous plot points and story ideas, and I can be assured that each and every one of them will pass the litmus test of plausibility, because I did the math up front.

This is a spiffy design for giant robot fans. Those titanic mecha arms will immediately grab the attention anybody who adores Jaegers.

Hello, and welcome back to Beyond NERVA! Today’s blog post is a special one, spurred on by the announcement recently about the Transport and Energy Module, Russia’s new nuclear electric space tug!

This is a fascinating system with a lot of promise, but has also gone through major changes in the last year that seem to have delayed the program. However, once it’s flight certified (which is to be in the 2030s), Roscosmos is planning on mass-producing the spacecraft for a variety of missions, including cislunar transport services and interplanetary mission power and propulsion.

Begun in 2009, the TEM is being developed by Energia on the spacecraft side and the Keldysh Center on the reactor side. This 1 MWe (4MWt) nuclear reactor will power a number of gridded ion engines for high-isp missions over the spacecraft’s expected 10-year mission life.

First publicly revealed in 2013 at the MAKS aerospace show, a new model last year showed significant changes, with additional reporting coming out in the last week indicating that more changes are on the horizon (there’s a section below on the current TEM status).

This is a rundown of the TEM, and its YaDEU reactor. I also did a longer analysis of the history of the TEM on my Patreon page (patreon.com/beyondnerva), including a year-by-year analysis of the developments and design changes. Consider becoming a Patron for only $1 a month for additional content like early blog access, extra blog posts and visuals, and more!

TEM Spacecraft

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The TEM is a nuclear electric spacecraft, designed around a gas-cooled high temperature reactor and a cluster of ion engines.

The TEM is designed to be delivered by either Proton or Angara rockets, although with the retirement of the Proton the only available launcher for it currently is the Angara-5.

Secondary Power System

Both versions of the TEM have had secondary folding photovoltaic power arrays. Solar panels are relatively commonly used for what’s known as “hotel load,” or the load used by instrumentation, sensors, and other, non-propulsion systems.

It is unclear if these feed into the common electrical bus of the spacecraft or form a secondary system. Both schemes are possible; if the power is run through a common electrical bus the system is simpler, but a second power distribution bus allows for greater redundancy in the spacecraft.

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The Primary propulsion system is the ID-500 gridded ion engine. For more information about gridded ion engines in general, check out my page on them here: https://beyondnerva.com/electric-propulsion/gridded-ion-thrusters/

The ID-500 was designed by the Keldysh Center specifically to be used on the TEM, in conjunction with YaEDU. Due to the very high power availability of the YaEDU, standard ion engines simply weren’t able to handle either the power input or the needed propellant flow rates, so a new design had to be come up with.

The ID-500 is a xenon-propelled ion engine, with each thruster having a maximum power level of about 35 kW, with a grid diameter of 500 mm. The initially tested design in 2014 (see references below) had a tungsten cathode, with an expected lifetime of 5000 hours, although additional improvements through the use of a carbon-carbon cathode were proposed which could increase the lifetime by a factor of 10 (more than 50,000 hours of operation).

Each ID-500 is designed to throttle from 375-750 mN of thrust, varying both propellant flow rate and ionization chamber pressure. The projected exhaust velocity of the engine is 70,000 m/s (7000 s isp), making it an attractive option for the types of orbit-altering, long duration missions that the TEM is expected to undertake.

The fact that this system uses a gridded ion thruster, rather than a Hall effect thruster (HET), is interesting, since HETs are the area that Soviet, then Russian, engineers and scientists have excelled at. The higher isp makes sense for a long-term tug, but with a system that seems that it could refuel, the isp-to-thrust trade-off is an interesting decision.

The initial design released at MAKS 2013 had a total of 16 ion thrusters on four foldable arms, but the latest version from MAKS-2019 has only five thrusters. The new design is visible below:

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The first design is ideal for the tug configuration: the distance between the thrusters and the payload ensure that a minimal amount of the propellant hits the payload, robbing the spacecraft of thrust, contaminating the spacecraft, and possibly building up a skin charge on the payload. The downside is that those arms, and their hinge system, cost mass and complexity.

The new design clusters only five (less than one third) thrusters clustered in the center-line of the spacecraft. This saves mass, but the decrease in the number of thrusters, and the fact that they’re placed in the exact location that the payload makes most sense to attach, has me curious about what the mission profile for this initial TEM is.

It is unclear if the thrusters are the same design.

Lovtsov, A.S., and Selivanov, M. Y. “FIRE TESTS OF HIGH POWER ION ENGINE FOR PERSPECTIVE TRANSPORT MODULES” 2014 http://naukarus.com/ognevye-ispytaniya-ionnogo-dvigatelya-vysokoy-moschnosti-dlya-perspektivnyh-transportnyh-moduley

Thermal Management

This may be the most interesting thing in in the TEM: the heat rejection system.

Most of the time, spacecraft use what are commonly called “bladed tubular radiators.” These are tubes which carry coolant after it reaches its maximum temperature. Welded to the tube are plates, which do two things: it increases the surface area of the tube (with the better conductivity of metal compared to most fluids this means that the heat can be further distributed than the diameter of the pipe) and it protects the pipe from debris impacts. However, there are limitations in how much heat can be rejected by this type of radiator: the pipes, and joints between pipes, have definite thermal limits, with the joins often being the weakest part in folding radiators.

The TEM has the option of using a panel-type radiator, in fact there’s many renderings of the spacecraft using this type of radiator, such as this one:

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However, many more renderings present a far more exciting possibility: a liquid droplet radiator, called a “drip refrigerator” in Russian. This design uses a spray of droplets in place of the panels of the radiator. This increases the surface area greatly, and therefore allows far more heat to be rejected. In addition it can reduce the mass of the system significantly, both due to the increased surface area and also the potentially higher temperature, assuming the system can recapture the majority of its coolant.

Work has been done both on the ground and in space on this system. The Drop-2 test is being conducted on the ISS, and multiple papers were published on it. It began in 2014, and according to Roscosmos will continue until 2024. http://www.tsniimash.ru/science/scientific-experiments-onboard-the-is-rs/cnts/experiments/kaplya_2/

Here it is being installed:

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Here’s an image of the results:

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A patent for what is possibly the droplet collection system has also been registered in Russia: https://yandex.ru/patents/doc/RU2607685C1_20170110

This system was also tested on the ground throughout 2018 (https://ria.ru/20181029/1531649544.html?referrer_block=index_main_2), and appears to have passed all the vacuum chamber ground tests needed. Based on the reporting, more in-orbit tests will be needed, but with Drop-2 already on-station it may be possible to conduct these tests reasonably easily.

I have been unable to determine what the working fluid that would be used is, but anything with a sufficiently low vapor pressure to survive the vacuum of space and the right working fluid range can be used, from oils to liquid metals.

For more on this type of system, check out Winchell Chung’s incredible page on them at Atomic Rockets: http://www.projectrho.com/public_html/rocket/heatrad.php#liquidradiator I will also cover them in the future (possibly this fall, hopefully by next year) in my coverage of thermal management solutions.

Of all the technologies on this spacecraft, this has to be the one that I’m most excited about. Some reporting (http://trudymai.ru/upload/iblock/a26/teploobmen-izlcheniem-dispergirovannykh-potokov-teplonositeley-kosmicheskikh-letatelnykh-apparatov.pdf ) says that this radiator can hit between 0.12 and 0.2 kW/kg system specific power!

Reaction Control Systems

Nothing is known of the reaction control system for the TEM. A number of options are available and currently used in Russian systems, but it doesn’t seem that this part of the design has been discussed publicly.

Additional Equipment

The biggest noticeable change in the rest of the spacecraft is the change in the spine structure. The initial model and renders had a square cross section telescoping truss with an open triangular girder profile. The new version has a cylindrical truss structure, with a tetrahedral girder structure which almost looks like the same structure that chicken-wire uses. I’m certain that there’s a trade-off between mass and rigidity in this change, but what precisely it is is unclear due to the fact that we don’t have dimensions or materials for the two structures. The change in the cross-section also means that while the new design is likely stronger from all angles, it makes it harder to pack into the payload fairing of the launch vehicle.

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The TEM seems like it has gone through a major redesign in the last couple years. Because of this, it’s difficult to tell what other changes are going to be occurring with the spacecraft, especially if there’s a significant decrease in electrical power available.

It is safe to assume that the first version of the TEM will be more heavily instrumented than later versions, in order to support flight testing and problem-solving, but this is purely an assumption on my part. The reconfiguration of the spacecraft at MAKS-2019 does seem to indicate, at least for one spacecraft, the loss of the payload capability, but at this point it’s impossible to say.

YaEDU Architecture

The YaEDU is the reactor that will be used on the TEM spacecraft. Overall, with power conversion system, the power system will weigh about 6800 kg.

Reactor

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The reactor itself is a gas cooled, fast neutron spectrum, oxide fueled reactor, designed with an electrical output requirement rather than a thermal output requirement, oddly enough (choice in power conversion system changes the ratio of thermal to electrical power significantly, and as we’ll see it’s not set in stone yet) of 1 Mwe. This requires a thermal output of at least 4 MWt, although depending on power conversion efficiency it may be higher. Currently, though, the 4 MWt figure seems to be the baseline for the design. It is meant to have a ten year reactor lifetime.

This system has undergone many changes over its 11 year life, and due to the not-completely-clear nature of much of its development and architecture, there’s much about the system that we have conflicting or incomplete information on. Therefore, I’m going to be providing line-by-line references for the design details in these sections, and if you’ve got confirmable technical details on any part of this system, please comment below with your references!

Fuel

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The fuel for the reactor appears to be highly enriched uranium oxide, encased in a monocrystalline molybdenum clad. According to some reporting (https://habr.com/en/post/381701/ ), the total fuel mass is somewhere between 80-150 kg, depending on enrichment level. There have been some mentions of carbonitride fuel, which offers a higher fissile fuel density but is more thermally sensitive (although how much is unclear), but these have been only passing mentions.

The use of monocrystalline structures in nuclear reactors is something that the Russians have been investigating and improving for decades, going all the way back to the Romashka reactor in the 1950s. The reason for this is simple: grain boundaries, or the places where different crystalline structures interact within a solid material, act as refractory points for neutrons, similarly to how a cracked pane of glass distorts the light coming through it through internal reflection and the disruption of light waves undergoing refraction in the material. There’s two ways around this: either make sure that there are no grain boundaries (the Russian method), or make it so that the entire structure – or as close to it as possible – are grain boundaries, called nanocrystalline materials (the preferred method of the US and other Western countries. While the monocrystalline option is better in many ways, since it makes an effectively transparent, homogeneous material, it’s difficult to grow large monocrystalline structures, and they can be quite fragile in certain materials and circumstances. This led the US and others to investigate the somewhat easier to execute, but more loss-intensive, nanocrystalline material paradigm. For astronuclear reactors, particularly ones with a relatively low keff (effective neutron multiplication rate, or how many neutrons the reactor has to work with), this monocrystalline approach makes sense, but I’ve been unable to find the keff of this reactor anywhere, so it may be quite high in theory.

It was reported by lenta.ru in 2014 (https://lenta.ru/news/2014/07/08/rosatom/ ) that the first fuel element (or TVEL in Russian) was assembled at Mashinostroitelny Zavod OJSC.

Reference was made (http://www.atomic-energy.ru/news/2015/07/01/58052 ) in 2015 to the fuel rods as “RUGBK” and “RUEG,” although the significance of this acronym is beyond me. If you’re familiar with it, please comment below!

In 2016, Dmitry Markov, the Director of the Institute of Reactor Materials in Zarechny, Sverdlovsk, reported that full size fuel elements had been successfully tested (https://xn--80aaxridipd.xn--p1ai/uchenye-iz-sverdlovskoj-oblasti-uspeshno-zavershili-ispytaniya-tvel-dlya-kosmicheskogo-yadernogo-dvigatelya/ ).

Coolant

The TEM uses a mix of helium and xenon as its primary coolant, a common choice for fast-spectrum reactors. Initial reporting indicated an inlet temperature of 1200K, with an outlet temperature of 1500K, although I haven’t been able to confirm this in any more recent sources. Molybdenum, tantalum, tungsten and niobium alloys are used for the primary coolant tubes.

Testing of the coolant loop took place at the MIR research reactor in NIIAR, in the city of Dimitrovgrad. Due to the high reactor temperature, a special test loop was built in 2013 to conduct the tests. Interestingly, other options, including liquid metal coolant, were considered (http://osnetdaily.com/2014/01/russia-advances-development-of-nuclear-powered-spacecraft/ ), but rejected due to lower efficiency and the promise of the initial He-Xe testing.

Power Conversion System

There have been two primary options proposed for the power conversion system of the TEM, and in many ways it seems to bounce back and forth between them: the Brayton cycle gas turbine and a thermionic power conversion system. The first offers far superior power conversion ratios, but is notoriously difficult to make into a working system for a high temperature astronuclear system; the second is a well-understood system that has been used through multiple iterations in flown Soviet astronuclear systems, and was demonstrated on the Buk, Topol, and Yenesiy reactors (the first two types flew, the third is the only astronuclear reactor to be flight-certified by both Russia and the US).

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In 2013, shortly after the design outline for the TEM was approved, the MAKS trade show had models of many components of the TEM, including a model of the Brayton system. At the time, the turbine was advertised to be a 250 kW system, meaning that four would have been used by the TEM to support YaEDU. This system was meant to operate at an inlet temperature of 1550K, with a rotational speed of 60,000 rpm and a turbine tip speed of 500 m/s. The design work was being primarily carried out at Keldysh Center.

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The Brayton system would include both DC/AC and AC/DC convertors, buffer batteries as part of a power conditioning system, and a secondary coolant system for both the power conversion system bearing lubricant and the batteries.

Building and testing of a prototype turbine began before the 2013 major announcement, and was carried out at Keldysh Center. (http://osnetdaily.com/2014/01/russia-advances-development-of-nuclear-powered-spacecraft/ )

As early as 2015, though, there were reports (https://habr.com/en/post/381701/ ) that RSC Energia, the spacecraft manufacturer, were considering going with a simpler power conversion system, a thermionic one. Thermionic power conversion heats a material, which emits electrons (thermions). These electrons pass through either a vacuum or certain types of exotic materials (called Cs-Rydberg matter) to deposit on another surface, creating a current.

This would reduce the power conversion efficiency, so would reduce the overall electric power available, but is a technology that the Russians have a long history with. These reactors were designed by the Arsenal Design Bureau, who apparently had designs for a large (300-500 kW) thermionic design. If you’d like to learn more about the history of thermionic reactors in the USSR and Russia, check out these posts:

https://beyondnerva.com/2018/11/08/nuclear-electric-propulsion-history-part-1-the-soviet-astronuclear-program/
https://beyondnerva.com/2019/03/10/topaz-international-part-1-enisy-the-soviet-years/
https://beyondnerva.com/2019/03/22/topaz-international-part-ii-the-transition-to-collaboration/

This was potentially confirmed just a few days ago by the website Atomic Energy (http://www.atomic-energy.ru/news/2020/01/28/100970 ) by the first deputy head of Roscosmos, Yuri Urlichich. If so, this is not only a major change, but a recent one. Assuming the reactor itself remains in the same configuration, this would be a departure from the historical precedent of Soviet designs, which used in-core thermionics (due to their radiation hardness) rather than out-of-core designs, which were investigated by the US for the SNAP-8 program (something we’ll cover in the future).

So, for now we wait and see what the system will be. If it is indeed the thermionic system, then system efficiency will drop significantly (from somewhere around 30-40% to about 10-15%), meaning that far less electrical power will be available for the TEM.

Radiation Shielding

The shielding for the YaEDU is a mix of high-hydrogen blocks, as well as structural components and boron-containing materials (http://www.atomic-energy.ru/news/2015/12/24/62211).

The hydrogen is useful to shield most types of radiation, but the inclusion of boron materials stops neutron radiation very effectively. This is important to minimize damage from neutron irradiation through both atomic displacement and neutron capture, and boron does a very good job of this.

Current TEM Status

Two Russian media articles came out within the past week about the TEM, which spurred me to write this article.

RIA, an official state media outlet, reported a couple days ago that the first flight of a test unit is scheduled for 2030. In addition:

Roscosmos announced the completion of the first project to create a unique space “tug” – a transport and energy module (TEM) – based on a megawatt-class nuclear power propulsion system (YaEDU), designed to transport goods in deep space, including the creation of long-term bases on the planets. A technical complex for the preparation of satellites with a nuclear tug is planned to be built at Vostochny Cosmodrome and put into operation in 2030. https://ria.ru/20200128/1563959168.html

A second report (http://www.atomic-energy.ru/news/2020/01/28/100970) said that the reactor was now using a thermionic power conversion system, which is consistent with the reports that Arsenal is now involved with the program. This is a major design change from the Brayton cycle option, however it’s one that could be considered not surprising: in the US, both Rankine and Brayton cycles have often been proposed for space reactors, only to have them replaced by thermoelectric power conversion systems. While the Russians have extensive thermoelectric experience, their experience in the more efficient thermionic systems is also quite extensive.

It appears that there’s a current tender for 525 million rubles for the TEM project by Roscosmos, according to the Russian government procurement website, through November 2021 (https://zakupki.gov.ru/epz/order/notice/ok504/view/common-info.html?regNumber=0995000000219000115 ), for

“Creation of theoretical and experimental and experimental backlogs to ensure the development of highly efficient rocket propulsion and power plants for promising rocket technology products, substantiation of their main directions (concepts) of innovative development, the formation of basic requirements, areas of rational use, design and rational level of parameters with development software and methodological support and guidance documents on the design and solution of problematic issues of creating a new generation of propulsion and power plants.”

Work continues on the Vostnochy Cosmodrome facilities, and the reporting still concludes that it will be completed by 2030, when the first mass-production TEMs are planned to be deployed.

According to Yuri Urlichich, deputy head of Roscosmos, the prototype for the power plant would be completed by 2025, and life testing on the reactor would be completed by 2030. This is the second major delay in the program, and may indicate that there’s a massive redesign of the reactor. If the system has been converted to thermionic power, it would explain both the delay and the redesign of the spacecraft, but it’s not clear if this is the reason.

For now, we just have to wait and see. It still appears that the TEM is a major goal of both Roscosmos and Rosatom, but it is also becoming apparent that there have been challenges with the program.

Conclusions and Author Commentary

It deserves reiterating: I’m some random person on the Internet for all intents and purposes, but my research record, as well as my care in reporting on developments with extensive documentation, is something that I think deserves paying attention to. So I’m gonna put my opinion on this spacecraft out there.

This is a fascinating possibility. As I’ve commented on Twitter, the capabilities of this spacecraft are invaluable. Decommissioning satellites is… complicated. The so-called “graveyard orbits,” or those above geosynchronous where you park satellites to die, are growing crowded. Satellites break early in valuable orbits, and the operators, and the operating nations, are on the hook for dealing with that – except they can’t.

Additionally, while many low-cost launchers are available for low and mid Earth orbit launches, geostationary orbit is a whole different thing. The fact that India has a “Polar Satellite Launch Vehicle” (PSLV) and “Geostationary Satellite Launch Vehicle” (GSLV) classification for two very different satellites drives this home within a national space launch architecture.

The ability to contract whatever operator runs TEM missions (I’m guessing Roscosmos, but I may be wrong) with an orbital path post-booster-cutoff, and specify a new orbital pat, and have what is effectively an external, orbital-class stage come and move the satellite into a final orbit is… unprecedented. The idea of an inter-orbital tug is one that’s been proposed since the 1960s, before electric propulsion was practical. If this works the way that the design specs are put at, this literally rewrites the way mission planning can be done for any satellite operator who’s willing to take advantage of it in cislunar space (most obviously, military and intelligence customers outside Russia won’t be willing to take advantage of it).

The other thing to consider in cislunar space is decommissioning satellites: dragging things into a low enough orbit that they’ll burn up from GEO is costly in mass, and assumes that the propulsion and guidance, navigation, and control systems survive to the end of the satellite’s mission. As a satellite operator, and a host nation to that satellite with all the treaty obligations the OST requires the nation to take on, being able to drag defunct satellites out of orbit is incredibly valuable. The TEM can deliver one satellite and drag another into a disposal orbit on the way back. To paraphrase a wonderful character from Sir Terry Pratchett (Harry King)“They pay me to take it away, and they pay me to buy it after.” In this case, it’s opposite: they pay me to take it out, they pay me to take it back. Especially in graveyard orbit challenge mitigation, this is a potentially golden opportunity financially for the TEM operator: every mm/s of mission dV can potentially be operationally profitable. This is potentially the only system I’ve ever seen that can actually say that.

More than that, depending on payload restrictions for TEM cargoes, interplanetary missions can gain significant delta-vee from using this spacecraft. It may even be possible, should mass production actually take place, that it may be possible to purchase the end of life (or more) dV of a TEM during decommissioning (something I’ve never seen discussed) to boost an interplanetary mission without having to pay the launch mass penalty for the Earth’s escape velocity. The spacecraft was proposed for Mars crewed mission propulsion for the first half of its existence, so it has the capability, but just as SpaceX Starship interplanetary missions require SpaceX to lose a Starship, the same applies here, and it’s got to be worth the while of the (in this case interplanetary) launch provider to lose the spacecraft to get them to agree to it.

This is an exciting spacecraft, and one that I want to know more about. If you’re familiar with technical details about either the spacecraft or the reactor that I haven’t covered, please either comment or contact me via email at beyondnerva@gmail.com

We’ll continue with our coverage of fluid fueled NTRs in the next post. These systems offer many advantages over both traditional, solid core NTRs and electrically propelled spacecraft such as the TEM, and making the details more available is something I’ve greatly enjoyed. We’ll finish up liquid fueled NTRS, followed by vapor fuels, then closed and open fueled gas core NTRs, probably by the end of the summer

If you’re able to support my efforts to continue to make these sorts of posts possible, consider becoming a Patron at patreon.com/beyondnerva. My supporters help me cover systems like this, and also make sure that this sort of research isn’t lost, forgotten, or unavailable to people who come into the field after programs have ended.

Light Freighter for intraorbital service between space colonies and industrial platforms, designed for the System States Era of my Orion’s Arm future history setting.

A timeline for my future history is to be found here: Timeline

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In the System States Era asteroid mining operations thrive throughout the asteroid belt and among the moons of Jupiter and Saturn the Martian terraforming program has left legacy: a sprawling archipelago of island stations and industrialized moons, Bernal Sphere's and O'Neill Cylinders, Spindle and Wheel cities, and a population of humanity growing into the millions. Space colonies are independent city-states and trade is their lifeblood. Entire generations are born and live their lives in spinning cylinders, bubbles, and torus shaped habitats, harvesting, mining, and fabricating all they need from the environment of the outer solar system.

Orion and Medusa style nuclear pulse freighters haul payloads of raw materials across interplanetary distances, while nuclear orbital transfer vehicles (OTV’s) provide light freight and passenger service between space habitats in Jupiter and Saturn orbit.

For a table of Delta V required for travel using Hohmann orbits among the moons of Saturn see
Why Saturn on Winchell Chung’s Atomic Rockets site. Scroll a little further down the page and you will find a Synodic Periods and Transit Times for Hohmann Travel table for Moons of Saturn.

Nuclear propulsion Systems: Operational Constraints

The abundance of various chemical ices for use as reaction mass among the moons of the outer system gas giants makes NERVA an excellent option for commercial application. Nuclear thermal rockets provide excellent efficiency; they also impose certain operational restrictions. The engine emits significant levels of radiation while firing and even after shut-down, and while passengers and crew are protected by the engines shadow-shield and hydrogen tanks, you wouldn’t want to point the engine at other spacecraft or space platforms. During the U.S. nuclear thermal rocket engine development program NFSD contractors had recommended that no piloted spacecraft approach to within 100 miles behind or to the sides of an operating NERVA I engine. The only safe approach to a spacecraft with a NERVA engine is through the conical “safe-zone” within the radiation shadow created by its shadow-shield and hydrogen tanks. Docking NERVA propelled spacecraft to a space station or habitat is problematic because structures protruding outside the conical safe-zone can reflect radiation back at the spacecraft, irradiating the passengers and crew.

These facts impose a set of mandatory operational parameters and flight rules for nuclear operation. An exclusion zone for nuclear propulsion (60 kilometers minimum) is imposed around every orbital platform. Orbital Guard units would hold broad discretionary powers—violate an exclusion-zone or disregard traffic-control and the local guard will cheerfully vaporize your spacecraft. No warning shots, no second chances. A crew that violates flight rules doesn’t live long enough to worry about fines or attorney fees, and the public’s time and funds are not wasted with trials of incompetent captains and crew.

Nuclear Freighters “park” propulsion modules in station-keeping orbit with their destination, and the freight/passenger module undocks, separating from its nuclear propulsion module, proceeding to birthing under thrust of a chemical maneuvering unit.

Because the nuclear propulsion modules are valuable, and are potentially deadly missiles if mishandled — codes to access the autonomous flight computer and possession of the nuclear propulsion module are temporarily handed over to the local orbital-guard for safe keeping.

For a good example of Space traffic control see the entry on Winchell Chung’s Atomic Rockets site here and scroll down to quote from Manna by Lee Correy.

At this point in my future history, 750 years post Martian colonization, spacecraft are essentially stacks of common modules which can be swapped out to suit application.

Independent Operators, like today’s truckers, might “own” only the CMOD (Command Module) with other units being leased per flight. The Freight Carrying Structural Spine, essentially a rigid frame with mountings for cargo modules, might be leased by the shipper and loaded with cargo (but owned by a separate freight transport supplier) and since different payloads mass differently it might be the responsibility of the shipper to lease suitable nuclear and chemical propulsion modules rated to the task. Passenger transport services might likewise lease passenger modules of varying capacity and Transport Brokerage firms would coordinate freight and passenger payloads assigned to same destinations and offer these in an open-bid market.

Propulsion Modules

Different payload masses require different propulsion module configurations, the light freighter detailed here requires only a single Solid-Core nuclear thermal rocket. A heavy payload freighter might use clusters of solid-core, or Open-Cycle Gas-Core, nuclear thermal rockets.

Design

3D models are my own conception based on various real-world proposals.

As research for the passenger/crew module I studied the POTV (Personnel Orbital Transfer Vehicle) pages 86-96 from NASA Technical Memorandum 58238 Satellite Power System: Concept Development and Evaluation Program Volume VI1 -Space Transportation available: here.

Propulsion for my light freighter is a Solid-Core NERVA Derivative, details available here.

In conversation Winchell Chung suggested the modification Cascade-Vanes: details available here.

Related Image: NUCLEAR OTV COMMERCIAL TRANSPORT

REFERENCE TUG
Engine	RL-10A-4-2 Chemical
@Thrust	99,100 N
@Mass Flow	21.87 kg/sec
Num Engines	x4
Total Thrust	396,400 N
Total Mass Flow	87.48 kg/sec
Specific impulse	462 sec
Exhaust Velocity	4,530 m/s
@Engine Mass	167 kg
Total Engine Mass	668 kg
Propellant	LOX/LH2
Thrust Power	6.5kW
Inert Mass	4,950 kg
Propellant Mass	62,000 kg

  Throughout the series I referred several times to a reference design for a cargo tug. That was put together using what might kindly be described as rectal numbers, assuming a tankage factor of 6% and 1.5 tons of remaining structure.

 I've gone back and run a preliminary estimation using a more detailed approach. In the process I created a spreadsheet that will allow users to enter their own values if desired. Errors are likely so use at your own risk; make a copy if you want to make changes.

 My little tug ended up just under 5 tons dry mass, 6.5kW power, 62 tons fuel capacity. The initial estimate was pretty close.

 I assumed the use of RL-10B-2 engines with modifications to allow unlimited restarts (Mr. Wolfe changed this to the RL10A-4-2, because the RL-10B-2 has too much mass). Most of the mass estimation relationships came from a U of Maryland slide series with some references from TU Delft and http://www.braeunig.us/. My estimates for cryocoolers are primitive at best, but the mass and power values are for a real off-the-shelf unit and should be reasonably accurate for cis-lunar space in the 50-100 ton fueled mass range. I also adjusted the thrust structure value to account for strength in two dimensions plus lunar landing gear; a note in that cell explains how to omit either option.

 Fuel hauling is assumed to use the craft's main propellant tanks. Cargo is assumed to be symmetrically attached to the thrust structure with low center of gravity. For cargo that is not being landed on the moon, attachment via a docking ring on the nose of the ship is also an option if a single bulky object is being transported.

 This more precise model comes to 4.95 tons of dry mass and 62 tons of fuel. All payloads assume 5% dV margins. It can land a 20-ton payload on the surface of the Moon and return it. It can deliver 50 tons of cargo to low Mars orbit and return empty. A one-way trip from LEO delivers 39 tons to EML1 or 46 tons to EML2.

Missions

Mission	Payload
Land payload on Lunar surface and return it	20,000 kg
Deliver payload to low Mars orbit and return empty	50,000 kg
One-way to EML1	39,000 kg
One-way to EML2	46,000 kg

 Fuel for a trip from EML1/2 takes two lunar surface fueling trips. Fuel for a trip from LEO takes three fueling trips from EML1/2 (requiring six surface trips). This is determined by engine Isp and the mass relationships used in the model, so the same is true if the vehicle is scaled up or down with similar thrust to weight ratio.

 I included three alternate configurations. The most relevant is sized to fit a standard Falcon payload fairing; 2.06t dry mass, 24.6t fuel capacity and would be launched engine-up empty on a 9 or fueled on a reusable heavy. Two much larger craft (120t and 500t fuel capacity) are presented as examples; the large craft could be launched on an Ariane V, Proton M or SLS block 1 with diameter as the main constraint. This would be a reasonable propellant depot in its own right. The huge craft would require an SLS block 2 due to both diameter and mass, plus a substantial supply of guts, stupidity or cash.

 Improvements to the model would be a better look at masses for a sunshield, cryocooler, payload hardpoints and docking system, as well as a frame with thermal/fragmentation protection between engines. A discrete tank model instead of a mass estimate based on volume is something I've done in another sheet and may migrate to this one in the future. A further set of refinements would be a look at ULA's integrated vehicle fluid management tech including gaseous H2/O2 lines for RCS thruster packs, combustion engine for reserve power and an electrolysis unit for converting waste water or payload water into additional fuel. I'd like to see a table or graph of total dV from surface to EML1/2 for a range of thrust to weight ratios as well. Lastly, some kind of mass estimate for a reusable (non-ablative) heatshield for the trip to LEO or LMO would be useful. I have not included estimates for any trip with aerobraking so the sheet would have to be reworked to account for that capability.

ELECTRIC TUG
Engines	Electromagnetic PIT
Specific Impulse	6,000 sec
Exhaust Velocity	58,900 m/s
Thrust	40 N
Power Requirement	1.6 MW
Power Plant	Solar Electric
Propellant	Water
Mass	33,400 kg

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

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

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

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

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

Estimating thrust requirements for average acceleration

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

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

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

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

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

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

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

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

Electric tug design

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

 Let's assume an electric thruster at Isp = 6000 and mission dV of 7000 m/s. Also assume a one-way trip (meaning fuel is available at both endpoints). Power alpha is assumed to be 18 kg/kW, whether that be nuclear or long-life solar. Thrusters will be the NuPIT design shown in the last post, using the design values for the 5 N, 200 kW unit at 2.75 kg/kW (550 kg per thruster).

 As a first guess let's try eight thrusters, 1.6 MW. That's 33.2 tons, for a dry mass without tanks of 73.2 tons. We will need approximately 10 tons of liquid water propellant; using a tankage fraction of 2% would be reasonable in this case, so tack on 200kg for tanks for a total dry mass of 73.4 tons. Actual propellant load is 9,273 kg, so tankage is sufficient. The half-fueled mass is 78,037 kg and approximate average acceleration is 0.513 mm/s. We're not even close. Trip time would be 157.9 days, or 2.3 one-way trips per year.

 Maybe 20 thusters / 4 MW? Power alpha would improve to about 15, yielding 71 tons of power and propulsion. 40 tons of payload and perhaps 0.4 tons of tankage gives a dry mass of 111.4 tons, fuel mass of 14.1 t and average acceleration of 0.844 mm/s. This is clearly not going our way. Trip time would be 106.3 days, or 3.4 one-way trips per year.

 Let's aim much higher, 50 thrusters / 10 MW. Power alpha would continue to improve to about 14, yielding 167.5 t of power and propulsion. 40 tons of payload and 0.6 tons of tankage gives a dry mass of 208.1 tons, 26.29 tons of fuel and an average acceleration of 1.13 mm/s. Trip time would be 71.7 days or 5.1 one-way trips per year.

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

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

Putting the tug to work

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

  So, a lunar ISRU plan would still start with a single chemical tug / lander as described in part 1. Using performance for the detailed reference tug, a 15-ton package can be delivered from LEO direct to the lunar surface. This will be 12.4 tons of ISRU equipment and 2.6 tons of spares (2.1 year supply). Refilling tug 1 will take 6 months, after which it can deliver 33 tons to EML1.

  In the meantime, an electric tug (call it tug A) will deliver a 9-ton fuel depot (135 ton capacity) to EML1. Let's use our 40 N / 1.6 MW / 33.4t / 6000 Isp vehicle from above. It does the job in about 92 days, which means there is a window of three months after the launch of the first ISRU package to get the 33.4t tug, 5.36t propellant and 9t payload into LEO.

  At the first lunar launch, 33 tons are delivered to the EML1 depot. Tug A will collect this and head to LEO, taking 127 days and consuming 7.45t of water. During this trip a LEO depot is launched, identical to the one at EML1. The tug turns back around and heads for EML1, taking 4.22t of water for the return trip and leaving 21.33t of cargo in LEO. This could be a mix of surface samples and water as desired. Let's assume five tons are samples and the rest is fuel.

  The return trip takes 72 days, during which tug 1 will have delivered another 33 tons to the EML1 depot. Tug A repeats its performance, returning to LEO with a full load of 23.78t water and another 5t of samples. At this point we are at 598 days elapsed since start of ISRU operations, which should be enough time to settle on and construct additional hardware to expand the lunar surface capacity.

  This is significantly longer than the all-chemical scenario and has an IMLEO of 129.71 tons, within a few tons of all-chemical. Hardware costs are higher since more of the mass is spacecraft and much less of it is fuel. The main benefit is that schedule pressures are greatly reduced; final design, construction and testing of the second round of ISRU plant is allowed more than a year and a half of time rather than two months. More operational data is available and the tolerance for mistakes or inefficiencies is higher. Another benefit is that this profile includes depots in LEO and at EML1; even if things do not progress beyond the first ISRU package the infrastructure is still useful for this and future projects.

  This baseline hardware could continue to deliver 21 tons of cargo to LEO every ~200 days for about a decade, eventually reaching 426 tons over 20 trips at a cost of 141 tons of Earth mass or a leverage of about 3 to 1. Things improve if we continue to expand, since about 44% of that mass was fuel to get the first ISRU plant in position; additional ISRU hardware is delivered using lunar propellant.

  The next phase would be to send more ISRU hardware. Tug A can pick up 33 tons at EML1, deliver 19.18t of net payload to LEO over 127 days, pick up a 17-ton package and head to EML1 in 109 days. All of the required propellant is lunar and picked up at EML1. Round trip time is 236 days (a bit under 8 months). The harvesting process run by tug 1 has a shorter turnover time of 6 months, so on average an extra 19 tons is accumulated at the depot. That's not quite enough to provide for a cargo landing, so tug A may not always be bringing a full load of cargo to LEO (meaning shorter round trips in practice).

  An alternative might be to use 12-ton packages that will fit into a Falcon 9 for cheaper launch costs; the delivery time for that is 98 days. If less cargo is returned to LEO then that trip time can be shortened as well; for example, 6 tons of return cargo plus round-trip fuel would make each leg of the trip take 98 days, or 196 days round-trip. Each 6.5-month trip would deliver another 10 tons of ISRU with two years of spares. Two electric tugs could deliver 80 tons of ISRU capacity in 26 months, roughly a single Mars synodic period. That would place 95 tons of ISRU with expected output of 950 tons of propellant annually at a cost of 9.5 tons of spares. Net propellant delivered to EML1 would be 505 tons annually, or could be 168 tons to LEO annually with chemical tugs. The annual demand for spares (both ISRU and depots) can be met in a single hardware run with minimal fuel costs, leaving 3/4 of the electric tug schedule open for assignments like delivering new chemical tugs or GEO debris retrieval (a mission that avoids the majority of the radiation belts and prolongs the tug's useful life).

 The total phase 2 IMLEO would be 134.4 tons, all of it hardware. Lunar mass to LEO during this period would only be another 48 tons since capacity is focused on buildup. This phase would run for 26 months, or a total of 46 months since first launch.

 Ongoing maintenance would require approximately 12 tons per year. The initial depots would be insufficient, so we need another 27t of hardware for fuel storage. If we rate all flight and depot hardware with a 10-year lifespan and pro-rate the replacement mass then we need an additional 12.2 tons annually (24.2t total). Depending on how the output is allocated, this could be considered an ongoing leverage of 23.5 tons in EML1 per ton IMLEO or 7.8 tons in LEO per ton IMLEO. Another way to look at it would be as a fuel supply for three manned Mars missions covering four synodic periods (104 months), or a full ISRU program length of 150 months (12.5 years). Overall Earth mass to LEO is then 511.15t to harvest 5,501.7 tons of gross lunar propellant, yielding 2,924.6 tons net lunar propellant at EML1. That's a gear ratio of about 5.7 to 1. If you are only interested in delivering fuel to LEO then you can net 972.9 tons, still a favorable 1.9 to 1 mass ratio. 511 tons is a lot of mass to launch, but only three payloads require a heavy lift vehicle: the initial chemical tug stack (62t fuel and 22t hardware, split across two Falcon H) and the two electric tugs (33.4t each, also requiring a Falcon H unless they can be built in parts and flown on two Vulcan launches). The remaining 360 tons would be delivered by 30 Falcon 9 launches, or by some combination of any price-competitive launchers with at least 12 tons of payload.

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

• 

      We suited up and cycled through the Can’s (Jupiter space station) main lock. The vehicle bay is just outside the lock, but the bay isn’t a particular room you can point at—it’s simply a big open space in the hollow part of the Can. All the small-sized vehicles are kept there and secured with elastic tie lines, to be sure they don’t bang into each other. All along the inner face of the Can are slots for berthing; when a vehicle needs to be fueled or worked over, it is pulled into a berth. Otherwise it’s moored a good distance from the Can’s skin, in high vacuum that does it no harm.

     Jenny and I clipped on to the mooring lines and pushed off. After a moment of coasting I turned so my feet pointed toward the shuttle and squirted my attitude jets. That slowed me to a crawl and I undipped from the line just as the shuttle swelled up to block my view of the opposite inner wall of the Can.

     I swung around, found a pipe and attached my own suit tie line to it. The shuttles are all different: each one was thrown together with whatever spare parts came to hand. The Roadhog—I’d silently christened it the moment my glove touched the pipe—looked like a conglomerate of castoffs until you studied the structure.

     It was a bit like an automobile chassis, all bones and no skin. The pilot was belted into a couch at the center. He was surrounded by pipes and struts and fuel tanks, without really obscuring his view. A small ion engine was mounted behind him. The whole thing was lumpy but balanced; spacecraft have to be stable.

     I glided over to the pilot’s couch and perched on top of the back rest. Around us, never closer than twenty yards, were other craft. A few had their running lights on; they were being checked over or preparing to go out. A big tubeshaped cargo hauler was moored right above us, shutting out the view through the bore of the Can. Below I could see stars, one of them a sharp, fierce blue diamond.

     I heard a faint clank as Jenny bumped into the shuttle. She secured her suit safety line and came swarming over to me. She touched helmets. “You know how to use the air tanks on this one, don’t you?”

     “Sure,” I said.
     “Take us over there, then,” she said, pointing to Berth G.

     I buckled myself into the pilot couch and reached out gingerly for the controls. You don’t use an ion engine inside the Can’s bay, or even nearby if you can help it. The backwash can knock a man head over heels a hundred meters away or snarl mooring lines. So I gently thumbed in the override on the shuttle’s air tanks, switched them over to the pipe system that led to the little maneuvering jets at the rear, and reached for the release button.

     “Forgotten anything?” Jenny said.
     “Huh?”
     “Our mooring lines.”

     “Oh.” I felt my face go red. I unbuckled and glided around the eight corners of the Roadhog, unhooking the elastic lines. They’re on retrieval coils, so as soon as I let go a line it retracted toward the Can walls.

     I sat back down in the couch. “All cleared, Captain.”

     She didn’t say anything. I carefully bled a little air into the pipes and felt a satisfying tug as we got underway. I gave us little bursts of air to maneuver around the cargo hauler overhead, and cut in the gyros to keep us from tumbling.

     We inched our way across the bay. I got back into the practice of looking three different directions at once; my neck started to ache. Human beings are built for navigating in two dimensions; our eyes are set in a line parallel to the ground. Outer space takes some getting used to. Even after you’ve trained your stomach to stop pushing the panic button when you’re in no-gee conditions, you have to keep reminding yourself that up and down are just as important as sideways. The adjustment is never perfect, because you’re trying to learn a set of reflexes our bodies just weren’t programmed to take. That’s why nogee maneuvering takes a lot of energy—you’re fighting yourself, all the way, whether you know it or not. I suppose that’s why kids like me are a little better at no-gee work and don’t tire so fast; our reflexes aren’t totally “set” yet.

     Berth G was a square-mouthed tube with bright lights lining the inside. I edged the Roadhog into the slot and brought us to a stop nearly perfectly; we couldn’t have been moving faster than a foot per second when we bumped into the buffer pads at the end.

     Jenny patted me on the shoulder and bounded away to fasten mooring lines.

     I felt good. I had proved that I could still handle a shuttle craft, despite being out of practice. And most of all, I was out in space again. It had been too long.

     That was the high point. The next five hours were something less than gratifying. Jenny took me over the Roadhog inch by inch, making me learn every valve and meter and strut on the contraption. I had forgotten a lot; the rest I hadn’t known at all.

     She made me draw a flow chart for the air pipes, after letting me inspect the Roadhog for five minutes. I thought I’d figured it out. When she handed the clip board back to me, covered with red marks, I found out that I had gotten everything exactly backwards.

     I checked out everything. Gyros, radio, first aid, fuel feeds, hauling collars, repair kit, spare parts. search lights, electrical system, navigation, backup systems, vector integrator—you name it, I had to find it, see if it worked, explain how I would repair it if it didn’t, and relate it to all the other systems it meshed with.

---

     We coasted into Berth G, freed the lines and Jenny gracefully swung into the pilot’s couch. She called in to the bridge and had an updated flight plan transmitted into the shuttle computer’s memory. Then I took over. I touched the fuel and oxygen gauges meaningfully, to prove I didn’t trust anybody’s word about that—like any good pilot—and ran quickly through the standard checklist. Jenny sat on the flat bench next to the couch, buckled herself in and gave me the high sign.

     I backed us cautiously out of the berth and brought the nose up to point at the circle of stars overhead. I gave Roadhog one burst of air through the rear jets and we coasted for the top of the Can in one long, clean straight line.

     We glided by the black shapes of parked craft and past the curved skin of the Can. A few viewports passed—they glowed with soft light and at one a woman looked out on the bay, smiled and waved. About 150 meters away in the opposite direction was the stubby cylinder at the axis of the Can, where the squash court and some hydroponics sections were. Behind us was the Far Eye, a little white dome at the tip of the central cylinder.

     The circle of stars grew as we rose toward the top of the Can. We passed the Sagan; thick hoses sprouted from her water tanks and led into several sockets in the Can skin. Jupiter peeked over the rim and quickly grew as we cleared the top of the Can. It was a crescent; the Can was moving sunward in its orbit.

     The shuttle shifted and murmured under me. The computer program was taking over. I punched the release button on the small control board and instantly felt a slight thrust. The ion engine had cut in. It made no noise; it’s a low-impulse system.

     We went straight up, away from the Lab, as though the Can was a cannon and we had been shot out of it. I was looking at Jupiter through the spaces in the Roadhog s floor.

     “What’s the radiation level like during the storm?”

     “High. Higher than they’ve ever seen before, Astrophysics says. Why, worried?”

     “Yup. I’m too young to be broiled in an electron shower. Are the shielding fields on?” I looked at my control panel.

     “Yes, they went on automatically when we left the Lab. Don’t worry.”

     “Don’t mind me, I’m a natural worrier.” I looked around at the superconducting bars that ring the Roadhog, though of course you can’t see the magnetic fields they produce. Those bars were all that kept Jupiter’s Van Allen belts from frying us alive.

     Radiation is a subtle thing. You can’t see it or taste it, but those little electrons and protons are better killers than a tiger. They are why the Lab wasn’t orbited in close to Jupiter, within the orbit of Io.

     Earth and Jupiter have one big thing in common: radiation belts. A man named Van Allen discovered them back in the early Space Age, around Earth. A little later Jupiter turned out to have them, too. Mars doesn’t, or Venus, or Mercury. Reason: no magnetic fields. Earth and Jupiter generate big magnetic fields around them, and those fields trap high-energy particles that the Sun throws out.

     They’re called belts because that’s what they look like—big donuts around Jupiter and Earth, seven or eight planetary radii in diameter. The Lab had to be located out beyond that donut or we’d be fried with radiation.

     Even as far from Jupiter as it is, the Can still has superconducting magnets all around the rim to deflect away charged particles—that’s what radiation is—because the outer edge of the Van Allen donut isn’t sharp, and a lot of deadly particles are still around, even that far from Jupiter.

     Why don’t Mars or Venus or Mercury have magnetic fields? My father tells me a planet has to do two things to produce them: first, have enough mass to squeeze its core until the metal inside melts. Second, rotate fast enough to churn the molten iron and nickel around. That forces electrical currents to flow, which makes a magnetic field. But Mars and Mercury just aren’t big enough to squeeze their cores. Venus is, but it hardly rotates at all. Which is a shame, really. Mars might have been a decent place to live if it had those fields. Van Allen belts are filled with charged particles that didn't hit the planet; the magnetic fields deflect them away from the atmosphere and into a sort of magnetic bottle that keeps them away from the planet’s surface.

     Earth has that; it’s safe for vulnerable human beings to walk around on. Jupiter has it; that’s another reason to think life can make a go of it there. Mars hasn’t, and Martian life never really had a chance; it was pelted from the start with every bit of garbage the sun throws out.

     The International Space Agency knew all this before they ever raised an expedition to Jupiter, of course; the Can is quite well equipped to deal with radiation. The Roadhog was a slightly different matter. It carried superconductors that could keep off medium-intensity radiation, but it couldn’t stop the big stuff that crops up during radiation storms. The Can, after all, gets a lot of help from sheer mass—the Lab water tanks line the outside of the Can and stop incoming particles before they can reach the living quarters.

     The Roadhog hasn’t got that mass. It’s a shuttle, engineered for speed and economy. So you don’t go out in it during radiation storms.

     Extra mass might have stopped the pellet that killed Ishi. Maybe there was an argument for putting shielding around the shuttles. Magnetic Helds don’t affect pieces of rock, because the rock is electrically neutral; only encasing a shuttle in heavy walls would make it really safe.