These are various rocket engines trying to harness the awesome might of antimatter. While the fuel is about as potent as you can get, trying to actually use the stuff has many problems.

Generally your spacecraft has metric tons of propellant, and a few micrograms antimatter fuel. The exceptions are the antimatter beam-core and positron ablative engines.

Nanograms of antimatter fuel are injected into some matter. The energy release is used to heat the propellant, which flies out the exhaust nozzle to create thrust.

Antimatter rockets have analogous exhaust velocity limits to nuclear thermal rockets. The higher the engine heat, the higher the exhaust velocity, which is a good thing. Unfortunately once the heat level reaches the liquefaction point, the engine melts. Which is a bad thing. This limits the maximum exhaust velocity.


(ed note: according to Adam Cowl, this equation applies when drives are power limited, based on the endurance of the engine rather than the energy of the fuel. I am not sure which engines beside antimatter qualify. Adam Cowl says the maximum mass ratio would be ~4.42, while the article below imples it will be ~4.9)

To those rocket engineers inured to the inevitable rise in vehicle mass ratio with increasing mission difficulty, antimatter rockets provide relief. The mass ratio of an antimatter rocket for any mission is always less than 4.9:1 [Shepherd, 1952], and cost-optimized mass ratios are as low as 2:1 [Forward, 1985]. In an antimatter rocket, the source of the propulsion energy is separate from the reaction fluid. Thus, the rocket's total initial mass consists of the vehicle's empty mass, the reaction fluid's mass, and the energy source's mass, half of which is the mass of the antimatter. According to the standard rocket equation, the mass ratio is now (assuming mr » me)


Δv = change in vehicle velocity (m/s)
ve = rocket exhaust velocity (m/s)
mi = initial mass of the vehicle (kg)
mf = final mass of the vehicle (kg)
mv = empty mass of the vehicle (kg)
mr = mass of the reaction fluid (kg)
me = mass of the energy source (kg)

The kinetic energy (K.E.) in the expellant at exhaust velocity (ve) comes from converting the fuel's rest-mass energy into thrust with an energy efficiency (ηe):


K.E. = kinetic energy (kg·m2/s2)
c = speed of light (3 × 108 m/s)

Solving Eq. (11.14) for the reaction mass (mr), substituting into Eq. (11.13), and solving for the energy source's mass (me) produces

We can find the minimum antimatter required to do a mission with a given Δv. We set the derivative of Eq. (11.15) with respect to the exhaust velocity ve equal to zero, and solving (numerically) for the exhaust velocity:

Substituting Eq. (11.16) into Eq. (11.13), we find that, because the optimal exhaust velocity is proportional to the mission Δv, the vehicle mass ratio is a constant:

The reaction mass (mr) is 3.9 times the vehicle mass (mv), while the antimatter fuel mass is negligible. Amazingly enough, this constant mass ratio is independent of the efficiency (ηe) with which the antimatter energy is converted into kinetic energy of the exhaust. (If the antimatter engine has low efficiency, we will need more antimatter to heat the reaction mass to the best exhaust velocity. The amount of reaction mass needed remains constant.) If we can develop antimatter engines that can handle jets with the very high exhaust velocities Eq. (11.16) implies, this constant mass ratio holds for all conceivable missions in the solar system. It starts to deviate significantly only for interstellar missions in which the mission Δv approaches the speed of light [Cassenti, 1984].

(ed note: Translation: to compensate for poor efficiency of antimatter energy converted into kinetic energy you do not need more reaction mass, you just need a few more milligrams of antimatter. Assuming the engine can resist being vaporized by the higher temperatures that come with the higher exhaust velocities.)

We can obtain the amount of antimatter needed for a specific mission by substituting Eq. (11.16) into Eq. (11.15) to get the mass of the energy source (me). The antimatter needed is just half of this mass. We find it to be a function of the square of the mission velocity (Δv) (essentially the mission energy), the empty vehicle's mass (mv), and the conversion efficiency (ηe):

(ed note: so the above equation is the important one, to figure how much antimatter fuel your spacecraft requires. Offhand I'd say the difference between antimatter solid core, beam core, and plasma core is the conversion efficiency (ηe) and the upper limit on antimatter per second fuel consumption set by the heat resistance of the engine)

The amount of antimatter calculated from Eq. (11.18) is typically measured in milligrams. Thus, no matter what the mission, the vehicle uses 3.9 tons of reaction mass for every ton of vehicle and an insignificant amount (by mass, not cost) of antimatter. Depending on the relative cost of antimatter and reaction mass after they have been boosted into space, missions trying to lower costs may use more antimatter than that given by Eq. (11.18) to heat the reaction mass to a higher exhaust velocity. If so, they would need less reaction mass to reach the same mission velocity. Such cost-optimized vehicles could have mass ratios closer to 2 than 4.9 [Forward, 1985].

The low mass ratio of antimatter rockets enables missions which are impossible using any other propulsion technique. For example, a reusable antimatter-powered vehicle using a single-stage-to-orbit has been designed [Pecchioli, 1988] with a dry mass of 11.3 tons, payload of 2.2 tons, and 22.5 tons of propellant, for a lift-off mass of 36 tons (mass ratio 2.7:1). This vehicle can put 2.2 tons of payload into GEO and bring back a similar 2.2 tons while using 10 milligrams of antimatter. Moving 5 tons of payload from low-Earth orbit to low Martian orbit with an 18-ton vehicle (mass ratio 3.6:1) requires only 4 milligrams of antimatter.

Antimatter rockets are a form of nuclear rocket. Although they do not emit many neutrons, they do emit large numbers of gamma rays and so require precautions concerning proper shielding and stand-off distance.

[Forward, 1985] Forward, Robert L., Brice N. Cassenti, and David Miller. 1985. Cost Comparison of Chemical and Antihydrogen Propulsion Systems for High AV Missions. AIAA Paper 85-1455, AIAA/SAE/ASME/ASEE 21st Joint Propulsion Conference, 8-10 July 1985, Monterey, California.

[Pecchioli, 1988] Pecchioli, M. and G. Vulpetti. 1988. A Multi-Megawatt Antimatter Engine Design Concept for Earth-Space and Interplanetary Unmanned Flights. Paper 88-264 presented at the 39th Congress of the International Astronautical Federation, Bangalore, India 8-15 October 1988.

[Shepherd, 1952] Shepherd, L. R. 1952. Interstellar Flight. Journal of the British Interplanetary Society. 11:149-167.

Antimatter Energy

Most of this is from Antiproton Annihilation Propulsion by Robert Forward.

From a practical standpoint, the proton-antiproton annihilation reaction produces two things: high-energy pions with an average kinetic energy of 250 MeV, and high-energy gamma rays with an average energy of 200 MeV.

Electron-positron annihilation just produces propulsion-worthless gamma rays, so nobody uses it for rockets. Except for the stranger antimatter engine designs.

To use the energy for propulsion, you have to either somehow direct the gamma rays and pions to shoot out the exhaust nozzle to produce thrust, or you have to used them to heat up a propellant and direct the hot propellant out the exhaust nozzle. To keep the crew and the computers alive you have to shield them from both gamma rays and pions. As far as the crew is concerned both reaction products come under the heading of "deadly radiation."

Charged Pions

Since pions are particles (unlike gamma rays) enough shielding will stop them all. Given an absorbing propellant or radiation shield of a specific density you can figure the thickness that will stop all the pions. This is the pion's "range" through that material.

In table 7-2 the columns under the yellow bar show how many centimeters (the "range") of the given stopping material is required to absorb 100 MeV of pion energy. The two sets of orange bars is because while the range is relatively constant for all high energies, the range becomes dramatically less at the point where the pion energy drops below 100 MeV (the "last 100 MeV").

For example: if the stopping material is water, absorbing 100 Mev from a 300 MeV hihg-energy pion requires 50 centimeters. But you only need 27 centimeters of water to absorb 100 MeV from a 75 MeV pion.

Since hydrogen, helium, and nitrogen have regrettably low densities the reaction chamber will have to operate at high pressure to get the density up to useful levels. "Useful" is defined as when the interaction range is shorter than the pion's mean life range. The Space Shuttle engines operated at a pressure of 213 atmospheres, 300 is a bit excessive. So of the gases nitrogen might be preferrable, even though you can get better specific impulse out of propellants with lower molecular weight.

Using detailed calculations they didn't explain, the report said hydrogen at 300 atm was about 65% efficient at converting the pion energy into heated propellant, while nitrogen at 100 atm was more like 95%.

Using more calculations that were not explained figure 7-4 was produced. The curve is the relative intensity of a charged pion at a given kinetic energy in MeV. The 125 MeV pions are the most intense (there are more of them), the average energy is 250 MeV.

Mean Life is the lifespan (not half-life) of a pion at that energy in nanoseconds. The range of a pion at that energy can be measured on the RANGE scales below, traveling through vacuum, hydrogen (H2) propellant at 300 atm, nitrogen (N2) propellant at 100 atm, and tungsten radiation shielding.

Gamma Rays

Sadly gamma rays cannot be used to propel the rocket (well, actually there are a couple of strange designs that do use gammas), all they do is kill anything living and destroy electronic equipment. So you have to shield the crew and electronics with radiation shielding. This is one of the big drawbacks to antimatter rockets. Gamma-rays would be useful if you were using antimatter as some sort of weapon instead of propulsion. But I digress.

A small number of "prompt" gamma-rays are produced directly from the annihilation reaction. The prompt gammas have a whopping 938 MeV, but they only contribute about 0.5% of the total. Almost ignorable.

A much larger amount of "delayed" gamma-rays are produced by the neutral pions decaying 90 attoseconds after the antimatter reaction. The spectrum peaks at about 70 MeV and trails off for many hundreds of MeV, with an average of 200 MeV.

Radiation Shielding

Most of this is from Antiproton Annihilation Propulsion by Robert Forward.

As mentioned above, the antimatter reaction is basically spitting out charged pions and gamma rays. The pions can be absorbed by the propellant and their energy utilized. The gamma rays on the other hand are just an inconvenient blast of deadly radiation traveling in all directions. The only redeeming feature is gamma rays are not neutrons, so at least they don't infect the ship structure with neutron embrittlement and turn the ship radioactive with neutron activation.

Since gamma rays are rays, not particles, they have that pesky exponential attenuation with shielding. It is like Zemo's paradox of Achilles and the tortoise, making the radiation shielding thicker reduces the amount of gamma rays penetrating but no matter how thick it becomes the gamma leakage never quite goes to zero. Particle shielding on the other hand have a thickness where nothing penetrates.

Gamma rays with energies higher than 100 MeV have a "attenuation coefficient" of about 0.1 cm2/g. Since tungsten has a density of 19.3 g/cm3 a tungsten radiation shield would have an attuation factor of 1.93 cm-1. Table 7-3 gives the attunation for various thickness of tungsten radiation shields.

This tells us that a 2 centimeter thick shield would absorb 97.9% of the gamma rays. 2.1×10-2 = 0.021 = 2.1%. 100% - 2.1% = 97.9%.

The main things that have to be shielded are the crew, the electronics, the cryogenic tankage, and the magnetic coils if this particular antimatter engine utilzes coils.

The radiation flux will be pretty bad. As an example, a ten metric ton rocket accelerating at 1 m/s2 will need a thrust level of 10,000 Newtons. If it has a specific impulse of 2000 s it will have an exhaust velocity of 20,000 m/s. This means the thrust power is Fp = (F * Ve ) / 2 = 100,000,000 watts = 100 megawatts.

Well, actually the report says 200 megawatts so obviously I made a mistake somewhere.

Anyway the thrust power basically is the fraction of the antimatter annihilation energy that becomes charged pions. Since 0.5% of the annihilation energy becomes prompt gamma rays, and the rest becomes 1.5 neutral pions (who become delayed gamma rays) and 3 charged pions then:

Eγ = (Eπ± * 1.506) - Eπ±


Eπ± = charged pion energy = thrust power
Eγ = gamma ray energy

So if the example rocket has 200 megawatts of thrust power, the gamma ray flux will be:

Eγ = (Eπ± * 1.506) - Eπ±

Eγ = (200 * 1.506) - 200

Eγ = 101.2 megawatts of lethal gamma rays

To shield the inanimate superconducting coils, table 7-3 tells us 10 centimeters of shield will give us an attenuation of 4.2×10-9, reducing the 101.2 megawatts down to 0.4 watts. The coil coolant systems should be able to handle that. The superconducting coils do not care about the biological dose since the coils are already dead.

But you do not get something for nothing. The 10 centimeters of coil shield prevent the radiation from hitting the coils but it does not make the radiation magically disappear. The coil shield will need a large heat radiator system capable of rejecting 101.2 megawatts of heat.

You will need more to shadow shield the living crew and sensitive electronics.

The report cites the American Institute of Physics handbook which mentions a 1 Curie source of gamma rays with an average energy of 100 MeV at a distance of 1 meter will expose you to 29 röntgen/hr (0.29 sievert per hour).

Our antimatter gamma rays have an average energy of twice that, 200 MeV not 100 MeV. So it becomes 58 röntgen/hr (0.58 sv/hour).

Let's assume the crew habitat module is 10 meters away from the engine instead of 1 meter. Radiation falls of according to the inverse square law. Inverse square of 10 times the distance is 1/102 or 1/100. So it becomes 58 / 100 = 0.58 röntgen/hr (0.0058 sv/hr).

That is the dose for a 1 Curie source. Our engine is much more radioactive than that.

Extrapolating further, a single 200 MeV gamma ray photon has 3.2×10-11 joules. This means a 101.2 megawatt source of 200 Mev gamma rays will produce 3×1018 gamma rays per second. This is equal to 8.5×107 Curies. Which is quite larger than 1 Curie.

1 Curie of 200 MeV gamma rays at a distance of 10 meters is 0.58 röntgen/hr. So 8.5×107 Curies will increase the dosage 8.5×107 times, to 4.9×107 röntgen/hr (490,000 sv/hr or 136 sv/second). This is very very bad since a mere 80 sieverts is enough to instantly put a person into a coma with certain death following in less than 24 hours. The poor crew will get that dose in about half a second. A shadow shield is indicated.

Looking at table 7-3 again, we see that 14 centimeters of tungsten has an attunation factor of 1.8×10-12. This will reduce the dose to 0.0000882 röntgen/hr (8.82×10-7 sv/hr) which the report describes as a reasonable dose for a space mission.

In the conceptual schematic, the reaction chamber is about 1 meter in diameter. The pressure walls have an equivalent thickness of 2 centimeters of tungsten, absorbing most of the gamma rays and coverting them into heat. The pressure walls are cooled by hydrogen flowing through channels in the wall. The hot hydrogen is sprayed as a film over the exhaust nozzle to protect it from the ultrahot hydrogen plasma blasting out from the antimatter reaction.

As per the calculations above, the superconducting coils are shielded with 10 centimeters of tungsten, with the thermal shields aimed at the antimatter annihilation point. 1 meter reaction chamber diameter plus 10 centimeters of shield makes the shield rings have a diameter of about 1.1 meter.

Also as per the calculations above, the personnel will be protected by a shadow shield 14 centimeters thick and 0.6 meters in diameter located 0.6 meters from the annihilation point. This will provide a 10 meter diameter shadow at a distance of 10 meters from the engine, for the habitat module and other ship parts to shelter in.

The reaction chamber is 2,200 kilograms, each thermal shield ring is 750 kilograms, and the shadow shield is 800 kilograms.

Solid Core

p-Nerva engine (NRX)
Thrust4.4×105 N
Thrust Power2.7 GW
Engine Mass11,000 kg
T/W >1.0yes
Specific Impulse1,100 sec to
1,300 sec
Exhaust Velocity10,790 m/s to
12,750 m/s
Fuelantiprotons (p)
Fuel Mass Flow13 μg/sec
(1.3×10-8 kg/s)
Mass Flow
40.7 kg/s
p-LH2 Mix1×10-6 kg p per
7,000 kg LH2
Borowski p engine
Thrust4.4×105 N
Thrust Power2.7 GW
Engine Mass7,000 kg
(less p containment)
Specific Impulse1,100 sec
Exhaust Velocity10,790 m/s
Fuelantiprotons (p)
Fuel Mass Flow15 μg/sec
(1.5×10-8 kg/s)
Mass Flow
41 kg/s

Basically a NERVA design where a tungsten antimatter target replaces the reactor.

A stream of antiprotons ( p ) antimatter fuel strike the tungsten target. The antiprotons annihilate protons inside the tungsten, producing gamma rays and pions. These are captured by the tungsten target, heating it. The tungsten target then heats the hydrogen propellant. Then the propellant rushes out the exhaust nozzle, creating rocket thrust.

Tungsten was chosen because it has an admirable effectiveness of stopping both the gamma rays and pions, a range of about 9 centimeters and a slowing down time of 0.5 nanoseconds. The tungsten is formed into a honeycomb, to allow the passage of propellant to be heated.

The tungsten also acts as the biological shadow shield.

Produces high thrust but the specific impulse is limited due to material constraints (translation: above a certain power level the blasted tungsten melts). Tungsten has a melting point of 2,683 K.

Predictably even though this engine has a thrust-to-weight ratio higher than one, the citizens are going to protest if you get the bright idea of using this rocket to boost payloads into orbit. Because an accident is going to be quite spectactular. You thought a nuclear explosion was bad, get a load of this!

According to Some Examples of Propulsion Applications Using Antimatter by Bruno Augenstein a tungsten block heated by antiprotons can heat hydrogen propellant up to a specific impulse of 1,000 to 1,300 seconds, depending up on the pressure the hydrogen operates at. This will require about one milligram (1×10-6 kg) of antiprotons per six or seven metric tons of hydrogen propellant, fed at a rate of 13 micrograms (1.3×10-8 kg) of antiprotons per second. One milligram of antiprotons has about the energy of an Aviation Thermobaric Bomb of Increased Power, or 43 tons of TNT.

According to Comparison of Fusion/Antiproton Propulsion Systems for Interplanetary Travel by Stanley K. Borowski (NASA Technical Memorandum 107030 AIAA–87–1814) a nuclear thermal engine needs all sorts of weird requirements to ensure nuclear criticality. Otherwise the reactor doesn't work. Antimatter, on the other hand, don't need no stinkin' criticality requirements. So the antimatter engine is much simpler.

All you need to do is make sure the tungsten target core is large enough to soak up most of the antimatter reaction products (so as to not waste antimatter energy and to protect the crew from radiation) and large enough to provide adequate hydrogen flow for cooling. Sadly there would be some neutron radiation due to positrons interacting with heavier nuclei. They figure the operating temperatures could be high enough to make the exhaust velocity around 9,810 m/s (Isp ~1,000 s). The tungsten core would be slightly smaller than a NTR reactor core, being a tungsten cylinder of about 80 cm diameter × 80 cm length. It would have a mass of 5,000 kg, assuming a 36% void fraction for the hydrogen coolant flow channels.

If you sized this engine for a crewed Mars mission, it would have a thrust of 4.4×105N, power level of 2.7 gigawatts, engine mass about 7,000 kg, and a specific impulse of about 1,100 sec (exhaust velocity of 10,790 m/s). Assuming a 100% deposition of antimatter energy in the tungsten and a 88.5% conversion efficiency into jet power, the engine would need a mass flow of 15 micrograms (1.5×10-8 kg) of antiprotons per second and a mass flow of 41 kg/sec of hydrogen propellant. For comparison a nuclear thermal rocket would need a burnup of about 33 milligrams (3.3×10-5) of U235 per second

Understand that the engine is going to require large masses of electric and magnetic field devices to safely store, extract, and inject the antiprotons into the tungsten without blowing the ship to tarnation. This is true of all antimatter powered rockets, but antimatter proponents tend to sweep this under the rug and seldom mention it in the weight estimates.

Gas Core

Gas-Core 5k sec
FuelAntiprotons (p)
Fuel Flow Rate2.25×10-8 kg/sec
Propellant Flow Rate0.9 kg/sec
Antimatter TargetTungsten
Annihilation Power4.05 GW
Thrust Power1.08 GW
Specific Impulse5,000 sec
Exhaust Velocity49,050 m/s
Thrust44,000 N
Cavity Radius1.2 m
Radiator Mass87,000 kg
Chamber Mass25,000 kg
Magnetic Coil Mass70,000 kg
Total Engine Mass182,000 kg
T/W Ratio2.5×10-2
Specific Power5.9 kW/kg
Gas-Core 1.25k sec
FuelAntiprotons (p)
Fuel Flow Rate1.125×10-8 kg/sec
Propellant Flow Rate14.3 kg/sec
Antimatter TargetTungsten
Annihilation Power2.025 GW
Thrust Power0.27 GW
Specific Impulse1,250 sec
Exhaust Velocity12,260 m/s
Thrust44,000 N
Cavity Radius1.2 m
Chamber Mass25,000 kg
Magnetic Coil Mass70,000 kg
Total Engine Mass95,000 kg
T/W Ratio4.7×10-2
Specific Power2.8 kW/kg

Antimatter rockets have analogous exhaust velocity limits to nuclear thermal rockets. Once the heat level reaches the liquefaction point (2,683 K), the tungsten core melts. This limits the solid core antimatter rocket's maximum exhaust velocity.

Rocket engineers quickly figured that if the antimatter rocket shared the same limitation as nuclear thermal rockets, perhaps they could use the same solutions. The nuclear thermal solution was the Gas Core NTR. May I present to you the Gas Core Antimatter Rocket. This is from Comparison of Fusion/Antiproton Propulsion Systems for Interplanetary Travel by Stanley K. Borowski.

The basic idea is to take the Gas Core NTR design, and replace the ball of fissioning uranium-235 gas with a ball of hot tungsten gas bombarded with a stream of antiprotons.

The tungsten gas will be a target for the antiprotons, being heated by the antimatter energy released, then heating up the hydrogen propellant by radiant heat. And because the tungsten is already vaporized, it can be safely heated to much higher that the 2,683 K which solid core antimatter engines are limited to.

Again, the task is easier because the GCNTR has to ensure the U235 gas is critical so as to undergo fission. Antimatter doesn't have to worry about that. For instance, the GCNTR requires a chamber pressure of 1,000 atmospheres to ensure the U235 achieves a critical mass. Antimatter version can get by on orders of magnitude less pressure. However, the antimatter version will require a tweek or two. Since the tungsten is vapor, an external magnetic field will be needed to trap the charged pions and follow-on decay products (the tungsten plasma can only capture 2/3 of the annihilation energy). The two candidate geometries for the magnetic field are Baseball Coil and Yin-yang. They will need a ferociously strong magnetic field, about 15 Tesla assuming the dimensions of the antimatter engine are about the same as the GCNTR.

Making some other assumptions based on the GCNTR, the report calculates that the antimatter power to be about 4.05 gigawatts, and require an antimatter flow rate of 22.5 μg/sec (2.25×10-8 kg/sec). This is with an assumed Isp of 5,000 sec, exhaust velocity of 49,050 m/s, propellant flow rate of 0.9 kg/sec, thrust of 44,000 newtons, and a propellant inlet temperature of 1,400 K.

If you do not do anything to capture the gamma-ray annihilation energy, it will hit the chamber walls and have to be removed as waste heat. 1.332 freaking gigawatts of the stuff (hydrogen regenerative cooling of the chamber walls remove an additional 0.018 GW). You'll need a heat radiator of about 193,000 kilograms (radiator specific mass of 19 kg/m2 and operating temperature of 1,225 K). 193 metric tons of heat radiator makes this propulsion system much less attractive. The heat radiator mass can be reduced to 87 metric tons if you raise the operating temperature to 1,500 K.

Alternatively you can alter some engine parameters to reduce the required antimatter fuel and antimatter power by half. Which also reduces the gamma ray waste heat by half. What you do is to increase the tungsten temperature to 3,250 K (the report is unclear as to what the value was before, something bigger than 2,683 K) and the propellant inlet temperature to the same. This drops the require antimatter power in half from 4.05 gigawatts to 2.025 GW and the antimatter flow rate from 22.5 μg/sec to 11.25 μg/sec. The waste gamma-ray annihilation energy drops from 1.332 GW to 0.675 GW. The propellant flow rate is drastically increased from 0.9 kg/sec to a whopping 14.3 kg/sec. This allows the propellant to absorb the 0.675 GW of gamma-ray energy, thus removing the need for the 87 metric tons of heat radiator.

The drawback is the increase in propellant flow rate catastrophically drops the specific impulse from 5,000 sec to a miserable 1,250 sec. Zounds! That is brutal. At that point you might as well use a fission gas-core NTR, it has a better specific impulse and the fuel is much cheaper.

Liquid Core Antimatter
Specific Impulse2,000 sec
Exhaust Velocity19,620 m/s
Thrust to Weight Ratio2.0
Specific Power190 kW/kg

Since the gas-core antimatter engine is either plagued by 87 metric tons of penalty weight or a catastrophic drop in specific impulse, engineers were wondering if the Liquid-core nuclear thermal rocket could be adapted to antimatter with better results.

In the fission version, a layer of liquid U235 is held to the spinning chamber walls by centrifugal force. Hydrogen propellant is injected through the chamber walls (cooling the walls), is heated by bubbling through the red-hot liquid uranium, emerges into the center of the chamber, and rushes with high velocity out the exhaust nozzle, creating thrust. Specific impulse between 1,300 to 1,500 seconds.

In the antimatter version, a 10 centimeter layer of red-hot liquid tungsten replaces the liquid uranium. It is sprayed with antiproton fuel to create annihilation energy. Since tungsten has a higher boiling point than uranium, at a chamber pressure of 10 atmospheres and an exhaust-to-chamber pressure ratio of 10-3, the antimatter liquid core could have a specific impulse up to 2,000 sec and an exhaust velocity of 19,620 m/s. Thrust-to-weight ratio about 2.0, specific power of 190 kW/kg. Which is better than the gas-core antimatter engine.

Forward Antimatter Gas Core
Exhaust Velocity24,500 m/s
Specific Impulse2,497 s
ReactorLiquid Core
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorMagnetic Nozzle

Robert Forward has an altenate gas core antimatter rocket. Microscopic amounts of antimatter are injected into large amounts of water or hydrogen propellant. The intense reaction flashes the propellant into plasma, which exits through the exhaust nozzle. Magnetic fields constrain the charged pions from the reaction so they heat the propellant, but uncharged pions escape and do not contribute any heating. Less efficient than AM-Solid core, but can achieve a higher specific impulse. For complicated reasons, a spacecraft optimized to use an antimatter propulsion system need never to have a mass ratio greater than 4.9, and may be as low as 2. No matter what the required delta V, the spacecraft requires a maximum of 3.9 tons of reaction mass per ton of dry mass, and a variable amount of antimatter measured in micrograms to grams.

Well, actually this is not true if the delta V required approaches the speed of light, but it works for normal interplanetary delta Vs. And the engine has to be able to handle the waste heat.

Plasma Core

AM: Plasma
980,000 m/s
99,898 s
Thrust61,000 N
29.9 GW
0.06 kg/s
17 kg/MW
AM: Plasma
7,840,000 m/s
799,185 s
Thrust49,000 N
0.2 TW
0.01 kg/s
RemassLiquid Hydrogen
3 kg/MW
AM: Plasma
Engine Mass
500,000 kg
ReactorPlasma Core
Thermal Accel:
Reaction Heat
Magnetic Nozzle

Similar to antimatter gas core, but more antimatter is used, raising the propellant temperature to levels that convert it into plasma. A magnetic bottle is required to contain the plasma.

Moderate Density
1016 atoms/cm3
1010 antiprotons/cm3 to
1012 antiprotons/cm3
7.6×10-7 N⋅s/cm3 to
9.8×10-6 N⋅s/cm3
45,000 m/s to
590,000 m/s
4,610 s to
60,000 s
High Density
1018 atoms/cm3
1012 antiprotons/cm3
8.1×10-5 N⋅s/cm3
49,000 m/s
4,950 s

LaPointe Antiproton Magnetically Confined Plasma Engine

In NASA report AIAA-89-2334 (1989) Michael LaPointe analyzes a pulsed antimatter rocket engine that confines neutral hydrogen gas propellant and antiprotons inside a magnetic bottle. Refer to the report if you want the actual equations

The hydrogen propellant is injected radially across magnetic field lines and the antiprotons are injected axially along magnetic field lines. The antimatter explodes, heating the propellant into plasma, for as long as the magnetic bottle can contain the explosion. After that, the magnetic mirror at one end is relaxed, forming a magnetic nozzle allowing the hot propellant plasma to exit. The cycle repeats for each pulse. Remember that the hydrogen nucleus is a single proton, convenient to be annihilated by a fuel antiproton.

The magnetic bottle contains the antiprotons, charged particles from the antimatter reaction, and the ionized hydrogen propellant. Otherwise all of these would wreck the engine. The magnetic bottle is created by a solenoid coil, with the open ends capped by magnetic mirrors.

LaPointe studied a range of densities for the hydrogen propellant.

At moderate to high densities the engine is a plasma core antimatter rocket. Compared to beam-core, the plasma core has a lower exhaust velocity but a higher thrust. The engine can shift gears to any desired exhaust velocity/thrust combination within its range by merely adjusting the amount of antiprotons and hydrogen gas injected with each pulse. And of course it can shift gears to any desired combination even outside its range by adding cold hydrogen propellant to the plasma (which is the standard method).

The reaction is confined to a magnetic bottle instead of a chamber constructed out of metal or other matter, because the energy of antimatter easily vaporizes matter.

At moderate hydrogen densities there is a problem with the hydrogen sucking up every single bit of the thermal energy, lots of the charged particle reaction products escapes the hydrogen propellant without heating up hydrogen atoms. This is a waste of expensive antimatter.

At high hydrogen densities there is a problem with bremsstrahlung radiation. Charged particles from the antimatter reaction create bremsstrahlung x-rays as they heat up the hydrogen. You want as much as possible of the expensive antimatter energy turned into heated hydrogen, but at the same time you don't want more x-rays than your engine (or crew) can cope with.

In the table, it does not list the thrust of the engine, instead it lists the "normalized" thrust. For instance the high density engine has a normalized thrust of 8.1×10-5 N⋅s/cm3. Don't panic, let me explain. You see, the actual thrust depends upon the volume of the magnetic bottle and the engine pulse rate (the delay between engine pulses). This lets you scale the engine up or down, to make it just the right size.

T = (Tnormalized / ΔT) * Bvol


T = thrust (Newtons)
Tnormalized = normalized thrust (N⋅s/cm3)
ΔT = pulse rate (seconds)
Bvol = volume of magnetic bottle (cm3)

Say your magnetic bottle had a radius of 1 meter (100 centimeters) and a height of 10 meters (1000 centimeters). Volume of a cylinder is V=πr2h, so the magnetic bottle has a volume of 3.14×107 cubic centimeters. A pulse rate of 10 milliseconds is 0.01 seconds. The high density engine has a normalized thrust of 8.1×10-5 N⋅s/cm3. What is the engine's thrust?

T = (Tnormalized / ΔT) * Bvol
T = (8.1×10-5 / 0.01) * 3.14×107
T = 0.0081 * 3.14×107
T = 254,340 Newtons

The propellant mass flow is:

mDotp = (mp * np * Bvol) / ΔT


mDotp = hydrogen propellant mass flow (kg)
mp = atomic mass of hydrogen (kg) = 1.672621777×10−27
np = hydrogen density (atoms/cm3)
Bvol = volume of magnetic bottle (cm3)
ΔT = pulse rate (seconds)

And obviously the antimatter mass flow is:

mDotp = (mp * np * Bvol) / ΔT


mDotp = antiproton fuel mass flow (kg)
mp = rest mass of antiproton (kg) = 1.672621777×10−27
np = antiproton density (antiproton/cm3)
Bvol = volume of magnetic bottle (cm3)
ΔT = pulse rate (seconds)

The optimum performance for LaPointe's engine was at a hydrogen propellant density of 1016 hydrogen atoms per cubic centimeters, and an antiproton density between 1010 and 1012 antiprotons per cubic centimeter. With an engine that can contain the reaction for 5 milliseconds (0.005 second), these densities produce a normalized thrust of 7.6x10-7 N⋅s/cm3 to 9.8x10-6 N⋅s/cm3 over a range of exhaust velocities (45,000 to 590,000 m/s). The propellant is only capturing about 2% of the antimatter heat, but at an acceptable level of bremsstrahlung x-rays.

The thrust can be increased by increasing the hydrogen propellant density to 1018cm-3, but then you start having problems with the hydrogen plasma radiatively cooling (losing its thrust energy). You'll have to expel the plasma no more than 200 or so μseconds (0.0002 second) after the antiprotons are injected. Assuming you can do that the engine will have a normalized thrust of 8.1×10-5 N⋅s/cm3 with an exhaust velocity of 49,000 m/s or so.

Key engineering issues:

  • Efficiently generating antiproton fuel on the ground (creating antimatter fuel is insanely expensive)
  • Antiproton containment (antimatter fuel tanks that won't blow up)
  • Designing strong enough magnetic field coils (magnetic field strong enough to contain hydrogen plasma created by exploding antimatter)
  • Switching system for efficient pulsed coil operation (allowing plasma to escape at precisely the right milisecond)
  • System to inject antiprotons into annihilation region (tranporting antimatter from the tank into the reaction chamber without any "accidents")
  • Radiation shielding (to protect the magnetic coils and the crew)

The superconducting magnetic coils will need not only radiation shielding from gamma rays created by the antimatter explosion, but also from the bremsstrahlung x-rays. The radiation shield will need to be heavy to stop the radiation, and extra shielding be needed to cope with to surface ablation and degradation. The majority of the engine mass will be due to radiation shielding, which will severely reduce the acceleration (drastically lowered thrust-to-weight ratio).

Antimatter Bottle

Antimatter Bottle
Antimatter Bottle
Exhaust Velocity78,480 m/s
Specific Impulse8,000 s
Thrust34,700 N
Thrust Power1.4 GW
Mass Flow0.44 kg/s
Total Engine Mass180,000 kg
Frozen Flow eff.80%
Thermal eff.85%
Total eff.68%
ReactorAntimatter Catalyzed
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorMagnetic Nozzle
Specific Power132 kg/MW

Antimatter fuel can be stored as levitated antihydrogen ice. By illuminating it with UV to drive off the positrons, a bit is electromagnetically extracted and sent to a magnetic bottle.

There it is collided with 60 g of heavy metal propellant (9 × 1024 atoms of lead or depleted uranium). Each antiproton annihilates a proton or neutron in the nucleus of a heavy atom. The use of heavy metals helps to suppress neutral pion and gamma ray production by reabsorption within the fissioning nucleus. If regolith is used instead of a heavy metal, the gamma flux is trebled requiring far more cooling.

A pulse of 5 μg of fuel (3 × 1018 antiprotons) contains 900 MJ of energy, and at a repetition rate of 0.8 Hz, a power level of 700 MWth is attained.

Compared to fusion, antimatter rockets need higher magnetic field strengths: 16 Tesla in the bottle and 50 Tesla in the throat. After 7 ms, this field is relaxed to allow the plasma to escape at 6 keV and 350 atm.

These high temperatures and pressures cause higher bremsstrahlung X-ray losses than fusion reactors. Furthermore, the antiproton reaction products are short-lived charged pions and muons, that must be exhausted quickly to prevent an increasing amount of reaction power lost to neutrinos. About a third of the reaction energy is X-rays and neutrons stopped as heat in the shields (partly recoverable in a Brayton cycle), another third escapes as neutrinos. Only the final third is charged fragments directly converted to thrust or electricity in a MHD nozzle.

D.L. Morgan, “Concepts for the Design of an Antimatter Annihilation Rocket,” J. British Interplanetary Soc. 35, 1982. (For use in this game, to keep the radiator mass within reasonable bounds, I reduced the pulse rate from 60 Hz to 0.8 Hz.)

Robert L. Forward, “Antiproton Annihilation Propulsion”, University of Dayton, 1985.

From High Frontier by Philip Eklund

Beam Core

AM: Beam
Exhaust Velocity100,000,000 m/s
Specific Impulse10,193,680 s
Thrust10,000,000 N
Thrust Power500.0 TW
Mass Flow0.10 kg/s
Total Engine Mass10,000 kg
ReactorAntimatter Catalyzed
Remass AccelAnnihilation
Thrust DirectorMagnetic Nozzle
Specific Power2.00e-05 kg/MW

Microscopic amounts of antimatter are reacted with equal amounts of matter. Remember: unless you are using only electron-positron antimatter annihilation, mixing matter and antimatter does NOT turn them into pure energy. Instead you get some energy, some charged particles, and some uncharged particles.

The charged pions from the reaction are used directly as thrust, instead of being used to heat a propellant. A magnetic nozzle channels them. Without a technological break-through, this is a very low thrust propulsion system.

All antimatter rockets produce dangerous amounts of gamma rays. The gamma rays and the pions can transmute engine components into radioactive isotopes. The higher the mass of the element transmuted, the longer lived it is as a radioisotope.

Positron Ablative

Positron Ablative
Exhaust velocity49,000 m/s

This engine produces thrust when thin layers of material in the nozzle are vaporized by positrons in tiny capsules surrounded by lead. The capsules are shot into the nozzle compartment many times per second. Once in the nozzle compartment, the positrons are allowed to interact with the capsule, releasing gamma rays. The lead absorbs the gamma rays and radiates lower-energy X-rays, which vaporize the nozzle material. This complication is necessary because X-rays are more efficiently absorbed by the nozzle material than gamma rays would be.

Drawbacks include the fact that you need 1836 positrons to equal the energy of a single anti-proton, and only half the positrons will hit the pusher plate limiting the efficiency to 50%.

This system is very similar to Antiproton-catalyzed microfission


Fusion propulsion uses the awesome might of nuclear fusion instead of nuclear fission or chemical power. They burn fusion fuels, and for reaction mass use either the fusion reaction products or cold propellant heated by the fusion energy.

Advantages include:

  • The exhaust velocity/specific impulse is attractively high

  • The fuel is so concentrated it is often measured in kilograms, instead of metric tons. Note this is not necessary true of the propellant.

Drawbacks include:

  • Mass flow/thrust is small and cannot be increased without lowering the exhaust velocity/specific impulse. And high exhaust velocity is one of the advantages of fusion propulsion in the first place.

  • The reaction is so hot that any physical reaction chamber would be instantly vaporized. So either magnetism or inertia is used instead, and those have limits.

  • The hot reaction will also vaporize the exhaust nozzle. So fusion propulsion tends to use exhaust nozzles composed of bladed laceworks and magnetism. These too have their limits.

  • Using open-cycle cooling to prevent the reaction chamber and nozzle from vaporizing also lowers the exhaust velocity/specific impulse.

  • Like fission propulsion, fusion produces lots of dangerous radiation.

There is a discussion of the problems with physical reaction chambers/exhaust nozzles here. There is a discussion of magnetic nozzles here.

Fusion Fuels

For more details about fusion fuels, go here.

Torchship Fusion

(ed note: Luke Campbell is giving advice to somebody trying to design a torchship. So when he says that magnetic confinement fusion won't work, he means won't work in a torchship. It will work just fine in a weak low-powered fusion drive.)

For one thing, forget muon catalyzed fusion. The temperature of the exhaust will not be high enough for torch ship like performance.

You might use a heavy ion beam driven inertial confinement fusion pulse drive, or a Z-pinch fusion pulse drive.

I don't think magnetic confinement fusion will work — you are dealing with a such high power levels I don't think you want to try confining this inside your spacecraft because it would melt.

D-T (deuterium-tritium) fusion is not very good for this purpose. You lose 80% of your energy to neutrons, which heat your spacecraft and don't provide propulsion. 80% of a terrawatt is an intensity of 800 gigawatts/(4 π r2) on your drive components at a distance of r from the fusion reaction zone. (see here for more about drive component spacing)

If we assume we need to keep the temperature of the drive machinery below 3000 K (to keep iron from melting, or diamond components from turning into graphite), you would need all non-expendable drive components to be located at least 120 meters away from the point where the fusion pulses go off.

(ed note: 120 meters = attunation 180,000. 800 gigawatts / 180,000 = 4.2 megawatts)

D-D (deuterium-deuterium) fusion gives you only 66% of the energy in neutrons. However, at the optimum temperature, you get radiation of bremsstrahlung x-rays equal to at least 30% of the fusion output power.

For a terawatt torch, this means you need to deal with 960 gigawatts of radiation. You need a 130 meter radius bell for your drive system to keep the temperature down.

(ed note: 130 meters = attunation 210,000. 960 gigawatts / 210,000 = 4.5 megawatts)

D-3He (deuterium-helium-3) fusion gives off maybe 5% of its energy as neutrons. A bigger worry is bremsstrahlung x-rays are also radiated accounting for at least 20% of the fusion output power. This lets you get away with a 66 meter radius bell for a terawatt torch.

(ed note: 66 meters = attunation 55,000. 250 gigawatts / 55,000 = 4.5 megawatts. I guess 4.5 megawatts is the level that will keep the drive machinery below 3000 k)

To minimize the amount of x-rays emitted, you need to run the reaction at 100 keV per particle, or 1.16 × 109 K. If it is hotter or colder, you get more x-rays radiated and more heat to deal with.

This puts your maximum exhaust velocity at 7,600,000 m/s, giving you a mass flow of propellant of 34.6 grams per second at 1 terawatt output, and a thrust of 263,000 Newtons per terawatt.

This could provide 1 G of acceleration to a spacecraft with a mass of at most 26,300 kg, or 26.3 metric tons. If we say we have a payload of 20 metric tons and the rest is propellant, you have 50 hours of acceleration at maximum thrust. Note that this is insufficient to run a 1 G brachistochrone. Burn at the beginning for a transfer orbit, then burn at the end to brake at your destination.

Note that thrust and rate of propellant flow scales linearly with drive power, while the required bell radius scales as the square root of the drive power. If you use active cooling, with fluid filled heat pipes pumping the heat away to radiators, you could reduce the size of the drive bell somewhat, maybe by a factor of two or three. Also note that the propellant mass flow is quite insufficient for open cycle cooling as you proposed in an earlier post in this thread.

Due to the nature of fusion torch drives, your small ships may be sitting on the end of a large volume drive assembly. The drive does not have to be solid — it could be a filigree of magnetic coils and beam directing machinery for the heavy ion beams, plus a fuel pellet gun. The ion beams zap the pellet from far away when it has drifted to the center of the drive assembly, and the magnetic fields direct the hot fusion plasma out the back for thrust.


Exhaust Velocity22,000 m/s
Specific Impulse2,243 s
Thrust108,000 N
Thrust Power1.2 GW
Mass Flow5 kg/s
Total Engine Mass10,000 kg
Specific Power8 kg/MW

Fuel: deuterium and tritium. Propellant: lithium. 1 atom of Deuterium fuses with 1 atom of Tritium to produce 17.6 MeV of energy. One gigawatt of power requires burning a mere 0.00297 grams of D-T fuel per second.

Note that Tritium has an exceedingly short half-life of 12.32 years. Use it or lose it. Most designs using Tritium included a blanket of Lithium to breed more fresh Tritium fuel.


H-B Fusion
Exhaust Velocity980,000 m/s
Specific Impulse99,898 s
Thrust61,000 N
Thrust Power29.9 GW
Mass Flow0.06 kg/s
Total Engine Mass300,000 kg
Specific Power10 kg/MW

Fuel is Hydrogen and Boron-11. Propellant is hydrogen. Bombard Boron-11 atoms with Protons (i.e., ionized Hydrogen) and you get a whopping 16 Mev of energy, three Alpha particles, and no deadly neutron radiation.

Well, sort of. Current research indicatates that there may be some neutrons. Paul Dietz says there are two nasty side reactions. One makes a Carbon-12 atom and a gamma ray, the other makes a Nitrogen-14 atom and a neutron. The first side reaction is quite a bit less likely than the desired reaction, but gamma rays are harmful and quite penetrating. The second side reaction occurs with secondary alpha particles before they are thermalized.

The Hydrogen - Boron reaction is sometimes termed "thermonuclear fission" as opposed to the more common "thermonuclear fusion".

A pity about the low thrust. The fusion drives in Larry Niven's "Known Space" novels probably have performance similar to H-B Fusion, but with millions of newtons of thrust.

It sounded too good to be true, so I asked "What's the catch?"

The catch is, you have to arrange for the protons to impact with 300 keV of energy, and even then the reaction cross section is fairly small. Shoot a 300 keV proton beam through a cloud of boron plasma, and most of the protons will just shoot right through. 300 keV proton beam against solid boron, and most will be stopped by successive collisions without reacting. Either way, you won't likely get enough energy from the few which fuse to pay for accelerating all the ones which didn't.

Now, a dense p-B plasma at a temperature of 300 keV is another matter. With everything bouncing around at about the right energy, sooner or later everything will fuse. But containing such a dense, hot plasma for any reasonable length of time, is well beyond the current state of the art. We're still working on 25 keV plasmas for D-T fusion.

If you could make it work with reasonable efficiency, you'd get on the order of ten gigawatt-hours of usable power per kilogram of fuel.

Professor N. Rostoker, et. al think they have the solution, utilizing colliding beams. Graduate Student Alex H.Y. Cheung is looking into turning this concept into a propulsion system.


He3-D Fusion
Exhaust Velocity7,840,000 m/s
Specific Impulse799,185 s
Thrust49,000 N
Thrust Power0.2 TW
Mass Flow0.01 kg/s
Total Engine Mass1,200,000 kg
Specific Power6 kg/MW

Fuel is helium3 and deuterium. Propellant is hydrogen. 1 atom of Deuterium fuses with 1 atom of Helium-3 to produce 18.35 MeV of energy. One gigawatt of power requires burning a mere 0.00285 grams of 3He-D fuel per second.


Fusion Containment

There are five general methods for confining plasmas long enough and hot enough for achieving a positive Q (more energy out of a reaction than you need to ignite it, "break even"):

  • Closed-field magnetic confinement
  • Open-field magnetic confinement
  • Inertial confinement (see D-D inertial fusion)
  • Electrostatic inertial confinement (see 6Li-H fusor)
  • Cold fusion (see H-B cat fusion)
  • Of these reactions, the fusion of deuterium and tritium (D-T), has the lowest ignition temperature (40 million degrees K, or 5.2 keV). However, 80% of its energy output is in highly energetic neutral particles (neutrons) that cannot be contained by magnetic fields or directed for thrust.

    In contrast, the 3He-D fusion reaction (ignition temperature = 30 keV) generates 77% of its energy in charged particles, resulting in substantial reduction of shielding and radiator mass. However, troublesome neutrons comprise a small part of its energy (4% at ion temperatures = 50 keV, due to a D-D side reaction), and moreover the energy density is 10 times less then D-T. Another disadvantage is that 3He is so rare that 240,000 tonnes of regolith scavenging would be needed to obtain a kilogram of it. (Alternatively, helium 3 can be scooped from the atmospheres of Jupiter or Saturn.)

    Deuterium, in contrast, is abundant and cheap. The fusion of deuterium to itself (D-D) occurs at too high a temperature (45 keV) and has too many neutrons (60%) to be of interest. However, the neutron energy output can be reduced to 40% by catalyzing this reaction to affect a 100% burn-up of its tritium and 3He by-products with D.

    The fusion of 10% hydrogen to 90% boron (using 11B, the most common isotope of boron, obtained by processing seawater or borax) has an even higher ignition temperature (200 keV) than 3He-D, and the energy density is smaller. Its advantage is that is suffers no side reactions and emits no neutrons, and hence the reactor components do not become radioactive.

    The 6Li-H reaction is similarly clean. However, both the H-B and 6Li-H reactions run hot, and thus ion-electron collisions in the plasma cause high bremsstrahllung x-ray losses to the reactor first wall.

    From High Frontier by Philip Eklund

    The samples below are from Nuclear Propulsion—A Vital Technology for the Exploration of Mars and the Planets Beyond (1987).

    There are two types of mission. One way missions go from planet A to planet B (AB or A→B) or from planet B to planet A (BA or B→A). Round trip (RT or A→A) missions go from A to B and back to A.

    The bottom line is that inertial confinement fusion is far superior to magnetic confinement fusion.

    Sample Closed-field
    Magnetic Confinement
    Fusion Rocket
    FuelD-3He (spin polarized)
    Specific Impulse20,000 s
    Mass Flow0.308 kg/s
    Engine Alpha5.75 kW/kg
    Engine Mass1,033,000 kg
    Payload Mass200,000 kg
    Fusion Rocket
    Specific Impulse270,000 s
    Mass Flow0.015 kg/s
    Engine Alpha110 kW/kg
    Engine Mass486,000 kg
    Payload Mass200,000 kg
    Sample Tokamak Fusion Rocket
    One-way continuous-burn constant-Isp trajectory


    DAB (A.U.)
    Mi (mT)
    Mp (mT)
    ML/Mi (%)
    τAB (days)
    ai (10-3 g0)
    Sample Tokamak Fusion Rocket
    Round-trip trajectory


    Sample Inertial Confinement Fusion Rocket
    Round-trip continuous-burn constant-Isp trajectory


    DAB (A.U.)

    ML/Mi (%)

    The above tables were calculated with the following equations:

    Wf = Mf * g0

    MB = Mf + MpB→A

    1 / α = Mi / MB

    1 / β = MB / Mf

    Pf = Mp / Mi

    RM = 1 / (α * β) (two way)

    RM = 1 / β (one way)

    τAB = (Isp / (F / Wf)) * (1 / β) * ((1 / α) -1) (equation 10)

    τBA = (Isp / (F / Wf)) * (1 / β - 1) (equation 11)

    τRT = τAB + τBA (equation 12a)

    τRT = (Isp / (F / Wf)) * (1 / (α * β) - 1) (equation 12b)

    DAB(m) = ((g0 * Isp2) / (F / Wf))) * (1 / β) * ((1 / sqrt(α)) - 1)2 (equation 13a)

    DBA(m) = ((g0 * Isp2) / (F / Wf))) * ((1 / sqrt(β)) - 1)2 (equation 14)

    DAB(m) = DBA(m) (equation 13b)


    αp = engine alpha (W/kg)
    DAB = distance between A and B (meters)
    DBA = distance between B and A (meters)
    Isp = engine specific impulse (seconds)
    IMEO = initial mass in Earth orbit (kg)
    MB = dry mass plus just propellant to travel from B to A (kg)
    ML = mass of payload (kg)
    MW = mass of engine (kg)
    Mf = dry mass (kg)
    Mi = initial mass in Earth orbit (kg)
    MpA→A = mass of propellant used traveling round-trip from A to B to A (kg)
    MpA→B = mass of propellant used traveling one-way from A to B (kg)
    MpB→A = mass of propellant used traveling one-way from B to A (kg)
    p = propellant mass flow (kg/s)
    Pf = propellant mass fraction
    RM = spacecraft mass ratio
    τAB = time to travel one way from A to B (seconds)
    τBA = time to travel one way from B to A (seconds)
    τRT = time to travel round trop from A to B to A (seconds)
    Wf = dry weight (Newtons)

    Inertial Confinement

    Inertial Confinement Fusion is in the Pulse section.

    Electrostatic Inertial

    H-Li6 Fusor Reactor
    H-Li6 Fusor
    Exhaust Velocity19,620 m/s
    Specific Impulse2,000 s
    Thrust67,100 N
    Thrust Power0.7 GW
    Mass Flow3 kg/s
    Total Engine Mass54,000 kg
    Frozen Flow eff.92%
    Thermal eff.90%
    Total eff.83%
    Remass AccelElectrostatic
    Thrust DirectorMagnetic Nozzle
    Specific Power82 kg/MW

    A Farnsworth-Bussard fusor is little more than two charged concentric spheres dangling in a vacuum chamber, producing fusion through inertial electrostatic confinement. Electrons are emitted from an outer shell (the cathode), and directed towards a central anode called the grid. The grid is a hollow sphere of wire mesh, with the elements magnetically-shielded so that the electrons do not strike them. Instead, they zip right on through, oscillating back and forth about the center, creating a deep electrostatic well to trap the ions of lithium 6 and hydrogen that form the fusion fuel. With a one meter diameter grid and a fuel consumption rate of 7 mg/sec, the fusion power produced is 360 MWth.

    Half of this energy is bremsstrahlung X-rays, which must be captured in a lithium heat engine. The other half are isotopes of helium (3He and 4He), each at about 8 MeV. (Overall efficiency is 36%). Since both products are doubly charged, a 4 MeV electric field will decelerate them and produce two electrons from each, producing an 18 amp current at extremely high voltage.

    An electron gun using this 4 million volt energy would emit electrons at relativistic speeds. This beam dissipates quickly in space, unless neutralized by positrons or converted into a free electron laser beam.

    “Inertia-Electrostatic-Fusion Propulsion Spectrum: Air-Breathing to Interstellar Flight,” R W. Bussard and L. W. Jameson, Journal of Propulsion and Power, v. 11, no. 2, pp. 365-372.

    (Philo Farnsworth, the farm boy who invented the television, spent his last years in a lonely quest to attain break-even fusion in his ultra-cheap fusor devices. His ideas are enjoying a renaissance, thanks to Dr. Bussard, and working fusion reactors are making an appearance in high school science fairs. On the theory that the fusor is power-limited, I have scaled down Bussard’s 10 GW design to 360 MW.)

    From High Frontier by Philip Eklund

    Magnetic Confinement

    Thrust Power200 GW
    Exhaust velocity8,000,000 m/s
    Thrust50,000 n
    Engine mass0.6 tonne
    T/W >1.0yes

    A magnetic bottle contains the fusion reaction. Very difficult to do. Researchers in this field say that containing fusion plasma in a magnetic bottle is like trying to support a large slab of gelatin with a web of rubber bands. Making a magnetic bottle which has a magnetic rocket exhaust nozzle is roughly 100 times more difficult.

    Since the engine is using a powerful but tightly controlled magnetic field, it might be almost impossible to have a cluster of several magnetic confinement fusion engines. The magnetic fields will interfere with each other.

    There are two main forms of magnetic bottles: linear (in a straight line) and toroidal (donut shaped, a linear bent into a circle with the ends joined together).

    Linear Fusion
    Gasdynamic Mirror
    Exhaust Velocity1,960,000 m/s
    Specific Impulse199,796 s
    Thrust47,000 N
    Thrust Power46.1 GW
    Mass Flow0.02 kg/s
    ReactorMagnetic Confinement
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle

    Also known as "Open-field magnetic confinement".

    Examples include the Gasdynamic Mirror, Hedrick Fusion Spacecraft, and the Santarius Fusion Rocket.

    3He-D Mirror Cell
    3He-D Mirror Cell
    Exhaust Velocity313,920 m/s
    Specific Impulse32,000 s
    Thrust10,600 N
    Thrust Power1.7 GW
    Mass Flow0.03 kg/s
    Total Engine Mass106,667 kg
    Frozen Flow eff.92%
    Thermal eff.90%
    Total eff.83%
    ReactorMagnetic Confinement
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Specific Power64 kg/MW

    Helium 3 is an isotope of helium, and deuterium (abbreviated D) is an isotope of hydrogen. The 3He-D fusion cycle is superior to the D-T cycle since almost all the fusion energy, rather than just 20%, is deposited in the plasma as fast charged particles.

    Magnetic containers with a linear rather than toroidal geometry, such as steady-state mirrors, have superior ratios of plasma pressure to magnet pressure (β >30%) and higher power densities necessary for reaching the high (50 keV) 3He-D operating temperatures.

    The mirror design shown is a tube of 11 Tesla superconducting magnetic coils, with choke coils for reflection at the ends. The magnets weigh 12 tonnes, plus another 24 tonnes for 60 cm of magnet radiation shielding and refrigeration. A mirror has low radiation losses (20% bremsstrahlung, 3% neutrons) compared to its end losses (77% fast charged particles). These losses limit the Q to about unity and prevent ignition. (This is not a problem for propulsion, since reaching break-even is not required to achieve thrust. The plasma is held in stable energy equilibrium by the constant injection of auxiliary microwave heating.)

    The Q can be improved by a tandem arrangement: stacking identical mirror cells end to end so that one’s loss is another’s gain. The exhaust exiting one end can be converted to power by direct conversion (MHD), and the other end’s exhaust can be expanded in a magnetic flux tube for thrust.

    Mirrors improved by vortex technology, called field-reversed mirrors, introduce an azimuthal electron current which creates a poloidal magnetic field component strong enough to reverse the polarity of the magnetic induction along the cylindrical axis. This creates a hot compact toroid that both plugs end losses and raises the temperature of the scrape-off plasma layer fourfold (to 2.5 keV), corresponding to a specific impulse of 32 ksec.

    Mirrors, like all magnetic fusion devices, can readily augment their thrust by open-cycle cooling.

    “Considerations for Steady-State FRC-Based Fusion Space Propulsion,” M.J. Schaffer, General Atomics Project 4437, Dec 2000.

    From High Frontier by Philip Eklund
    Toroidal Fusion

    Also known as "Closed-field magnetic confinement".

    Discovery II
    Discovery II
    MC Fusion
    Exhaust Velocity347,000 m/s
    Specific Impulse35,372 s
    Thrust18,000 N
    Thrust Power3.1 GW
    Mass Flow0.05 kg/s
    ReactorMagnetic Confinement
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Wet Mass1,690,000 kg
    Dry Mass883,000 kg
    Mass Ratio1.91 m/s
    ΔV225,258 m/s
    Specific Power3.5 kW/kg (3,540 W/kg)
    Initial Acceleration1.68 milli-g
    Payload172,000 kg
    Length240 m
    Diameter60 m wide
    D-T Fusion Tokamak
    D-T Fusion Tokamak
    Exhaust Velocity66,800 m/s
    Specific Impulse6,809 s
    Thrust66,800 N
    Thrust Power2.2 GW
    Mass Flow1 kg/s
    Total Engine Mass197,000 kg
    Frozen Flow eff.77%
    Thermal eff.85%
    Total eff.65%
    ReactorMagnetic Confinement
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Specific Power88 kg/MW

    Of all the fusion reactions, the easiest to attain is a mixture of the isotopes of hydrogen called deuterium and tritium (D-T). This reaction is “dirty”, only 20% of the reaction power is charged particles (alphas) that can be magnetically extracted with a diverter for power or thrust. The remaining energy (neutron, bremsstrahlung, and cyclotron radiation) must be captured in a surrounding jacket of cold dense Li plasma. The heated lithium is either exhausted as open-cycle coolant, or recirculated through a heat engine (to generate the power needed for the microwave plasma heater).

    The 2 GWth magnetically-confined reactor shown uses eight poloidal superconducting 30 Tesla coils, twisted into a Tokamak configuration. These weigh 22 tonnes with stiffeners and neutron shielding.

    The pulsed D-T plasma, containing tens of megamps, is super-heated by 50 MW of microwaves or colliding beams to 20 keV. The Q (gain factor) is 40. Closed field line devices such as this can ignite and burn, in which case the Q goes to infinity and microwave heating is no longer needed. However, since ignition is inherently unstable (once ignited, the plasma rapidly heats away from the ignition point), the reactor is kept at slightly below ignition.

    Fuel is replenished at 24 mg/sec by gas puffing to maintain a plasma ion density of 5 × 1020/m3 at 26 atm. At a power density of 250 MWth /m3, the lithium-cooled first wall has a neutron loading of 1 MW/m2 and a radiation loading of 5 MW/m2.

    More advanced vortex designs, which do away with the first wall, separate the hot fusion fuel from the cool lithium plasma by swirling the mixture. The thermal efficiency is 50% in open-cycle mode.

    Williams, Borowski, Dudzinski, and Juhasz, “A Spherical Torus Nuclear Fusion Reactor Space Propulsion Vehicle Concept for Fast Interplanetary Travel,” Lewis Research Center, 1998.

    (The Tokamak used in High Frontier is a smaller lower tech version of the Lewis design, which uses aneutronic 3He-D fuel.)

    From High Frontier by Philip Eklund
    Direct Fusion Drive
    Mass Schedule
    ItemmassTotal mass
    Enginesx21,357 kg2,714 kg
    Radiatorsx266.75 kg133.5 kg
    Magnetic Nozzlesx271.43 kg142.86 kg
    D2 tanksx13,432 kg3,432 kg
    He3 Tankx10.22 kg0.22 kg
    Insulating shellx1200 kg200 kg
    Optical trussx210 kg20 kg
    Gimbalsx23.75 kg7.5 kg
    Laser comnx22 kg4 kg
    Orbiter Payloadx1270 kg270 kg
    Landerx1230 kg230 kg
    TOTAL7,154 kg
    TypeDirect Fusion Drive
    Num Enginesx2
    Thrust5 N
    Total Thrust10 N
    Isp10,000 sec
    Exhaust Vel98,100 m/s
    Power1 MW
    Total Power2 MW
    Num Radiatorsx2
    Radiator Area50 m2
    Radiator Cap.840 kW
    Thrust16 N
    Specific Impulse12,500 s
    Exhaust Velocity122,600 m/s
    Thrust Power2 MW
    Radius0.3 m

    This is currently the front-runner for something that can actually be made into a real working fusion drive. The specific impulse is fantastic, but predictably the thrust is measured in hummingbird-powers. Regardless of the low thrust, it can deliver a metric ton of payload to Pluto in 3.75 years flat, instead of chemical rocket New Horizon's pathetic 30 kilograms taking freaking nine years.

    Even better is the fact that it can run bi-modal: switching from generating thrust into generating electrical power. This comes in real handy if the spacecraft is traveling further than the orbit of Mars from Sol, solar power becomes pretty worthless in the outer solar system .

    It is a magnetic-confined fusion design, also a afterburner engine, so you can shift gears: increasing thrust by lowering specific impulse.

    It also plans to use the practically aneuronic (almost no deadly neutron radiation) D-3He fusion fuel so the ship does not need a massive neutron-radiation shield, which would savagely cut into the allowed payload mass.

    Understand that D-3He is only "practically" aneuronic because some of the deuterium stubbornly refuses to fuse with 3He as it should but instead unhelpfully fuses with other deuterium in the fuel and thus creates neutron radiation. Worse: the D+D reaction produces tritium, and if the tritium accumulates it can fuse with more deuterium and create even more neutron radiation. The Direct Fusion Drive deals with this by selective 3He heating, and by doing its darnedest to flush any tritium created out the exhaust stream ASAP. This does not totally eliminate the neutron radiation but any reduction is a win.

    The drawback is the reaction is harder to ignite, and Helium-3 is scarce stuff (at least on Terra). The report suggest trying to scrape together the scraps of Helium-3 found naturally in helium-rich natural gas wells. Currently most commercial Helium-3 comes from the decay of tritium triggers in nuclear warheads. Tritium can be produced by bombarding lithium-6 with neutrons, then tritium has a half-life of 12.3 years, decaying into Helium-3. Forget about mining Luna for Helium-3, it isn't concentrated enough to be economical. The best extraterrestrial source is the atmosphere of Saturn, it exists in Jupiter's atmo but the gravity is so great it is hard for the harvest ships to escape.

    To deal with the fact that the D-3He fusion reaction is harder to ignite, the engines uses a innovative way of altering the fields. This raises the plasma temperature high enough to kindle the reaction.

    The Bremsstrahlung and synchrotron radiation emitted from the plasma heat up the engine components. The coolant which prevents the engine from melting is sent through a Brayton heat engine to run an electrical power generator. The coolant is then sent to heat radiators. The power is used to energize the electromagnetic coils, the radio-frequency plasma heater, and any other power needs the spacecraft has.


    Direct Fusion Drive (DFD) is a conceptual low radioactivity, nuclear-fusion engine designed to produce both thrust and electric power for interplanetary spacecraft. The concept is based on the Princeton field-reversed configuration reactor invented in 2002 by Samuel A. Cohen, and is being modeled and experimentally tested at Princeton Plasma Physics Laboratory, a US Department of Energy facility, and modeled and evaluated by Princeton Satellite Systems. As of 2018, the concept has moved on to Phase II to further advance the design.


    The Direct Fusion Drive (DFD) is a conceptual fusion-powered spacecraft engine, named for its ability to produce thrust from fusion without going through an intermediary electricity-generating step. The DFD uses a novel magnetic confinement and heating system, fueled with a mixture of helium-3 (He-3) and deuterium (D), to produce a high specific power, variable thrust and specific impulse, and a low-radiation spacecraft propulsion system. Fusion happens when atomic nuclei, comprising one species in a hot (100 keV or 1,120,000,00 K) plasma, a collection of electrically charged particles that includes electrons and ions, join (or fuse) together, releasing enormous amounts of energy. In the DFD system, the plasma is confined in a torus-like magnetic field inside of a linear solenoidal coil and is heated by a rotating magnetic field to fusion temperatures. Bremsstrahlung and synchrotron radiation emitted from the plasma are captured and converted to electricity for communications, spacecraft station-keeping, and maintaining the plasma's temperature. This design uses a specially shaped radio waves (RF) "antenna" to heat the plasma. The design also includes a rechargeable battery or a deuterium-oxygen auxiliary power unit to startup or restart DFD.

    The captured radiated energy heats to 1,500 K (1,230 °C; 2,240 °F) a He-Xe fluid that flows outside the plasma in a boron-containing structure. That energy is put through a closed-loop Brayton cycle generator to transform it into electricity for use in energizing the coils, powering the RF heater, charging the battery, communications, and station-keeping functions. Adding propellant to the edge plasma flow results in a variable thrust and specific impulse (in other words it can shift gears) when channeled and accelerated through a magnetic nozzle; this flow of momentum past the nozzle is predominantly carried by the ions as they expand through the magnetic nozzle and beyond, and thus, function as an ion thruster.


    The construction of the experimental research device and most of its early operations were funded by the US Department of Energy. The recent studies —Phase I and Phase II— are funded by the NASA Institute for Advanced Concepts (NIAC) program. A series of articles on the concept were published between 2001 and 2008; the first experimental results were reported in 2007. Numerous studies of spacecraft missions (Phase I) were published, beginning in 2012. In 2017 the team reported that "Studies of electron heating with this method have surpassed theoretical predictions, and experiments to measure ion heating in the second-generation machine are ongoing." As of 2018, the concept has moved on to Phase II to further advance the design. The full-size unit would measure approximately 2 m in diameter and 10 m long.

    Stephanie Thomas is vice president of Princeton Satellite Systems and also the Principal Investigator for the Direct Fusion Drive.

    Projected performance

    Analyses predict that the Direct Fusion Drive would produce between 5-10 Newtons thrust per each MW of generated fusion power, with a specific impulse (Isp) of about 10,000 seconds and 200 kW available as electrical power. Approximately 35% of the fusion power goes to thrust, 30% to electric power, 25% lost to heat, and 10% is recirculated for the RF heating.

    Modeling shows that this technology can potentially propel a spacecraft with a mass of about 1,000 kg (2,200 lb) to Pluto in 4 years. Since DFD provides power as well as propulsion in one integrated device, it would also provide as much as 2 MW of power to the payloads upon arrival, expanding options for instrument selection, laser/optical communications, and even transfer up to 50 kW of power from the orbiter to the lander through a laser beam operating at 1080 nm wavelength.

    The designers think that this technology can radically expand the science capability of planetary missions. This dual power/propulsion technology has been suggested to be used on a Pluto orbiter and lander mission, a well as integration on the Orion spacecraft to transport a crewed mission to Mars in a relatively short time (4 months instead of 9 with current technology).

    From the Wikipedia entry for DIRECT FUSION DRIVE

    I serve on the External Council of NIAC, NASA's Innovative Advanced Concepts program, an organization that funds and encourages innovative ideas that are applicable to the US space program. NIAC had a meeting this June in Washington, DC, and there I heard a presentation describing an innovation that I've awaited for many years: the promise of a rocket that is propelled by the energy of nuclear fusion. Chemical rockets, the workhorses of our present space program, produce too little energy per kilogram of fuel and have exhaust velocities and specific impulse values that are too small. Elaborate and very expensive aerospace engineering efforts are necessary to get around the inadequacies of chemical rockets. The high energies and high exhaust velocities provided by fusion for space propulsion are exciting prospects. In particular, fusion rockets should have specific impulse values (how much push you get per mass of fuel) that are 10 to 100 times larger than those of chemical rockets. Amazingly, there is now a line of development that may make fusion rockets possible. To begin this discussion, I want to review a few basic ideas about nuclear fusion.

    The energy source that heats most stars is the nuclear fusion of hydrogen. For our Sun, it is a multi-step process involving the strong and weak interactions that starts with four protons (symbol p, the nucleus of a hydrogen atom) and finishes with the production of helium-4 (4He, a mass-4 helium nucleus containing two neutrons and two protons), along with some neutrinos, positrons, and gamma rays. In the deep interior of the Sun the temperature and pressure are high enough to fuse two protons to deuterium (d, mass-2 hydrogen nucleus containing one neutron and one proton), to fuse deuterium and a proton to helium-3 (3He, mass-3 helium nucleus containing one neutron and two protons), and to fuse two helium-3 nuclei to one helium-4 (4He, two protons and two neutrons) plus two free protons. This is the so-called p-p fusion chain that dominates fusion processes in all stars having the mass of our Sun or smaller.

    Fusion machines on Earth like the large and expensive ITER tokomak machine being developed in France must operate at lower temperatures and pressures than those in a star and must attempt to produce usable energy from controlled fusion using a simpler single-step process that fuses deuterium and tritium (t, mass-3 hydrogen nucleus containing two neutrons and one proton) into a 4He and a neutron (reaction: d+t4He+n). Much of the energy from this process is given to the neutron with a kinetic energy of 14.1 million electron volts (14.1 MeV). This neutron steals valuable energy and presents hazards for both radiation-sensitive materials and for human operators. Usually the fast neutrons must be moderated with massive shielding and captured in a lithium blanket to produce additional energy. The break-even temperature, the temperature to which the d+t plasma (ionized atoms) must be heated in order to produce more fusion energy than was required to create it, is 13.6 thousand electron volts (13.6 keV). Fusion machines are designed to contain such plasmas in a strong magnetic field while the plasma is heated, fusion occurs, and energy extracted.

    The problem with applying such fusion-machine designs to fusion-driven space propulsion is that they are typically very large and expensive, they have not yet produced any useful fusion energy, they have no obvious exhaust ejection path for propulsion, and they would produce large quantities of 14.1 MeV neutrons in close proximity to payloads and crew, with their neutrons removing energy needed for propulsion and requiring many tons of shielding material. It is apparent that d+t fusion, while possibly useful for power applications on Earth, is inappropriate for space applications. Clearly, a different approach is needed.

    Such a new approach may be the direct fusion drive rocket (DFDR) design of the Princeton Plasma Physics Laboratory, which has provided the basis for a new NIAC Phase I grant to Stephanie Thomas of Princeton Satellite Systems. At the NIAC meeting I attended, she discussed a mission to Pluto using a DFDR drive. The DFDR geometry is a cylinder that uses the so-called "field-reversed configuration". This means that the loading of the plasma in the system employs a trick. A plasma is initially created within a magnetic field from pinch coils at the two ends of a cylindrical polycarbonate vacuum vessel. This field is aligned in along the axis of the cylinder and causes the ions to orbit around the lines of magnetic flux, preventing them from moving away from the axis of the cylinder and sustaining the field. Then the current in the pinch coils is abruptly reversed, applying a new external magnetic field in the opposite direction. This abrupt change compresses and heats the plasma and creates a layered and increased magnetic field, with flux lines in the inner region in one direction and those in the outer region in the opposite direction. The outer magnetic field lines attempt to expand outward but are confined within the cylinder by a set of passive warm-superconducting "flux conserver" pancake coils spaced along the length of the cylinder. These coils develop induced currents of thousands of amperes when field lines attempt to cross them, keeping the magnetic flux lines within the cylinder. The trapped ions execute figure-8 orbits at the boundary between the inner and outer fields.

    The DFDR is designed to fuse deuterium and helium-3 to form a proton and a helium-4 nucleus (reaction: d+3Hep+4He), liberating 18.3 MeV of usable charged-particle kinetic energy in each fusion reaction. Note that neutrons are not produced in this primary reaction. However, because deuterium is present in the plasma, the fusion of two deuterium nuclei to tritium and a proton or to helium-3 and a neutron with a kinetic energy of 2.45 MeV (reactions: d+dt+p or d+d3He+n) are both secondary reactions that will also occur in the heated plasma. Further, if tritium from the first reaction is allowed to accumulate, one could again face the problem of dealing with 14.1 MeV neutrons from d+t fusion. Fortunately, the DFDR design includes mechanisms for suppressing d+d reactions by selective 3He heating and for moving tritium from the plasma into the exhaust stream as it is produced.

    We note, however, that 3He is an expensive fuel, mainly a small byproduct derived from helium-rich natural gas wells. Its use in fusion requires a considerably higher plasma temperature for fusion. While the break-even temperature for d+t fusion is 13.6 keV, the break-even temperature for d+3He is 58 keV, more than three times higher. Thus, an effective method for heating the plasma to this high temperature is a key element of the DFDR design.

    This heating is accomplished by applying to the plasma an "odd-parity rotating magnetic field" produced by four rectangular figure-eight shaped antenna coils surrounding the cylinder above, below, and on the sides. The drive currents in the four antenna coils are sequenced so that the resulting magnetic field is always perpendicular to the cylinder axis and rotates at a drive frequency in the MHz region. The "odd parity" description means that the front and rear parts of the antenna coils have oppositely circulating currents and produce magnetic field that point in opposite directions. This drive field oscillates near the cyclotron resonance frequency of 3He and therefore selectively transfers more heat to 3He ions than to d ions. In particular, 3He ions are heated to about 100 keV while d ions are heated to only 50 keV. This differential heating enhances 3He+d fusion with respect to d+d fusion by a large factor, minimizing the production of neutrons and tritium.

    The use of the rotating magnetic field brings another advantage: following fusion ignition of the plasma, the same antennas can be used to directly extract electric power from the plasma ions through induction. The result is a fusion rocket that also directly generates its own electric power, a very significant advantage for space missions that venture far from the Sun, where solar panels are not efficient.

    As illustrations of the utility of the DFDR, Princeton Satellite Systems has designed and described in some detail a spectrum of space missions that might use the fusion rocket technology. These include:

    1. the transport of the James Webb Space Telescope from low-earth orbit to the desired Lagrange-2 halo orbit by using a 1 megawatt (1 MW) DFSR
    2. a Pluto Orbiter/Lander sample-return mission, the subject of the NIAC grant mentioned above, using a 2 MW DFDR (x2 DFDR, each with 5 N and 1 MW)
    3. the deflection of an Apophis-type asteroid away from an Earth-collision orbit using a reusable "space tug" with a 5 MW DFDR engine
    4. a manned mission to Mars with a 46 day one-way travel time using a cluster of eight DFDRs generating power totaling 160 MW and delivering 2,000 newtons of thrust (x8 DFDR, each with 250 N and 20 MW)
    All of these missions would be cheaper and faster than similar missions using chemical rockets. Clearly a mature DFDR technology would bring a "game changing" renaissance to our space program.

    Therefore, the questions are: how soon will the research on fusion rockets be completed, and when can we begin to use them? The present incarnation of this technology, the PRFC-2, is only a prototype unit that has demonstrated ion heating to a few kV with a 15 KW rotating magnetic field for a duration of 100 ms. The system is reported to be behaving as predicted by theoretical modeling and shows good confinement time and no unanticipated instabilities. The warm-superconductor flux conserver coils are behaving as expected and can sustain induced currents of up to 3,100 amperes.

    However, there is a long way to go from this test configuration to a usable fusion rocket. To achieve fusion temperatures, the rotating magnetic field drive must deliver around 200 KW. The detailed design of the magnetic nozzle that shapes and directs the fusion products in the exhaust is a work in progress. The expectations are that the work will ultimately result in fusion engines capable of producing power of 1 to 20 megawatts, thrusts of 5 to 250 newtons, and specific impulses of 5,000 to 50,000 seconds.

    The timeline provided by Princeton Satellite Systems and the Princeton Plasma Physics Laboratory shows the planned PFRC-3 test unit demonstrating fusion around 2019, the PFRC-4 unit demonstrating operational thrust and energy production on the ground around 2024, the first robotic space mission being operational in about 2030, and a Mars Orbital Mission around 2038. This is probably an optimistic schedule, and it critically depends on whether the Department of Energy and NASA will provide funding that supports these developments.

    All I can say is that we really need fusion rocket technology, the DFDR is the most promising fusion rocket design I have seen, and I hope the timeline can be maintained.


    Modular Aneutronic Fusion Engine,IAC-12,C4,7-C3.5,10

    Direct Fusion Drive Rocket for Asteroid Deflection, J. B. Mueller, et al.

    Direct Fusion Drive for a Human Mars Orbital Mission, M. Paluszek, et al.,

    Fusion-Enabled Pluto Orbiter and Lander

    From THE DIRECT FUSION DRIVE ROCKET by John G. Cramer (2016)
    Flow-Stabilized Z-Pinch
    Flow-Stabilized Z-Pinch
    Sample Thrusters
    Pinch Length10 m50 m
    Temp80 keV100 keV
    Current5 MA10 MA
    Input Power1.8×1012W8.4×1012W
    Fusion Power3.3×1012W9.9×1012W
    Mass Flow0.095 kg/s0.53 kg/s
    Exhaust Velocity3.5×106 m/s1.3×106 m/s
    Specific Impulse356,800 sec132,500 sec
    Thrust3.3×105 N6.8×105 N

    This is from Advanced Space Propulsion Based on the Flow-Stabilized Z-Pinch Fusion Concept (2006) and Sustained neutron production from a sheared-flow stabilized Z-pinch (2019)

    I'm not sure I believe these numbers, they are getting close to being a torchship. Not quite, the Saturn V first stage puts out a hundred times more thrust. Having said that, the Robert Werka FFRE engine has a comparable exhaust velocity of 5.2×106 m/s but its thrust is a pathetic 43 Newtons. Not kilo-Newtons, just Newtons. The Flow-Stabilized Z-Pinch cranks out 7,700 times as much thrust.

    Remember: zeta-pinch or z-pinch works by taking a stream of plasma and sending a bolt of electricity down the center axis. The Lorentz force crushes the plasma toward the center axis. If the plasma is made of fusion fuel and the crush is sufficiently strong, the fuel undergoes nuclear fusion. If the reaction chamber is a linear tube (instead of a torus as is sometimes used), you just have to have one end open and capped with a magnetic nozzle to make a fusion rocket engine.

    The trouble is that the plasma is wildly unstable, with several gross disruption modes. Blasted thing squirms at high velocity like an angry snake being stepped on. Specifically it writhes so fast that only miniscule fraction of the fuel gets a chance to undergo fusion. Once the plasma stream bends enough to hit the chamber walls the electricity shorts out, the electrical current is no longer down the plasma axis, and the Lorentz force vanishes. No more fusion.

    But in 2006 researchers at the University of Washington figured out how to stabilize the plasma stream using what is known in fluid dynamics as sheared axial flow. And in 2019 other researchers at UWash working with sheared flow managed to stablize a z-pinch reaction for five thousand times longer (16 μs) than an unstablized z-pinch. A neutron signature of nuclear fusion was detected for 5 μs, which was also five thousand times longer than an unstablized z-pinch can manage.

    Flow-Stabilized Z-Pinch Fusion Space Thruster

    To utilize sheared axial flow stabilization, you need some flow. Since most rocket engines make the propellant flow as their basic operating principle, sheared axial flow fits in like a hand in a glove. The plasma flow undergoes nuclear fusion, releasing far more energy than any mere chemical reaction can manage. All you need to do is make the far end of the reaction chamber open to act as a rocket nozzle.

    Since no external magnetic fields are required to confine the plasma, massive magnetic coils can be dispensed with. Which is a good thing, those heavy coils really cut into the payload mass. Z-pinch fusion rockets are typically much lower in mass than conventional magnetic-confinement fusion engines, which vastly improves their specific power rating.

    In the diagram above, fusion fuel is injected at left at velocity V. Electrical current leaves inner electrode, travels through the axis of the fuel flow, and leaves fuel at flared end of outer electrode (nozzle at right side). The axial current in fuel creates the Lorentz force (mag field) which pinches the fuel flow. The pinch causes the fuel to undergo thermonuclear fusion. This accelerates the fusion reaction products to velocity Ve (which is a heck of a lot faster than a chemical reaction can do). The products exit the nozzle, creating thrust.

    The outer electrode exit end is flared. This is to give the fusion reaction products a chance to cool down from star-core-hot temperatures. Otherwise the heat would vaporize the electrode and void the rocket engine's warranty.

    A small amount of the hot fusion products are bleed off the other end of the engine. They enter the direct energy converter, creating electrical power. This harvesting of energy is common with many fusion propulsion designs, since fusion engines tend to be power hogs. Some kind of alternate electrical power is used as a "pilot light" to start things off.

    In the Sample Thruster table, the power needed for the axial current is listed under "Input Power".

    The electrical power can be used in the axial current, to supply power to the spacecraft's habitat module, and/or to energize a magnetic nozzle. Such a nozzle can improve the conversion of fusing product thermal power into thrust.

    As with most fusion engines cold propellant can be injected in order to shift gears (raise the thrust at the cost of lowering the exhaust velocity/specific impulse).

    The Sample Thruster table shows two thrusters. Both were optimized for the fusion fuels they utilize. The equations used can be found in the first report, they are rather involved.

    Magnetio Inertial Confinement

    Magnetio Inertial Fusion is in the Pulse section.

    Fusion Engines

    To make the fusion reactor into a fusion rocket, the fusion energy has to be used to accelerate reaction mass. The method will determine the exhaust velocity/specific impulse, which is the critical variable in the delta V equation.

    There are three types of energy that come from fusion reactions:

    • Plasma thermal energy: When the fusion fuel undergoes fusion, the fuel atoms are ionized into useful hot plasma ions containing most of the fusion energy in a convenient easy-to-use form. We like plasma thermal energy.

    • Neutron energy: Many fusion reactions or side reactions also produce deadly and worthless neutron radiation. It is lethal to human beings. It can cause neutron embrittlement and neutron activation in the engine parts. Neutron energy is considered to be wasted energy.

    • Bremsstrahlung radiation energy: This occurs when the hot plasma ions from the fusion reaction collide with the electrons (which are there because "ionization of fusion fuel atoms" means "ripping off their electrons and tossing them into the plasma soup"). Bremsstrahlung steals the hot ion's useful plasma thermal energy and converts it into worthless and dangerous x-rays plus cold ions. This is also considered to be wasted energy.

    Pure fusion rockets use the fusion products themselves as reaction mass. Fusion afterburners and fusion dual-mode engines use the fusion energy (plasma thermal energy, neutron energy, and bremsstrahlung radiation energy) to heat additional reaction mass. So afterburners and dual-mode reduce the exhaust velocity in order to increase thrust.

    • Pure fusion rockets use just the plasma thermal energy, and just the fusion products as reaction mass. The neutron and bremsstrahlung radiation energy is considered to be waste.
      This mode has the highest exhaust velocity/specific impulse and the lowest thrust/propellant mass flow of the three fusion engine types.

    • Fusion afterburners use just the plasma thermal energy, but adds extra cold reaction mass to be heated by plasma energy. Again neutron and bremsstrahlung are wasted.

    • Dual-mode use the neutron and bremsstrahlung radiation energy to heat a blanket of cold reaction mass which thrusts out of separate conventional exhaust nozzles. In addition a Dual-mode can switch into Pure Fusion mode.
      This mode has the highest thrust/propellant mass flow and the lowest exhaust velocity/specific impulse.

    Dr. Stuhlinger notes that high-thrust mode allows fast human transport (but low payloads) while high-specific-impulse mode allows cargo vessels with large payload ratios (but long transit times). He compares these to sports cars and trucks, respectively.

    In the Santarius Fusion Rocket using D-3He fusion:

    Santarius Fusion Rocket
    D-3He Fusion
    ModeSpecific ImpulseThrust
    Pure Fusion1×106 sec88 N
    Afterburner5×105 sec to
    1×104 sec
    125 N to
    5,000 N
    Dual-Mode7×102 sec to
    7×101 sec
    12,500 N to
    125,000 N

    Pure Fusion Engines

    Pure fusion rockets use just the plasma thermal energy, and just the fusion products as reaction mass.

    The advantage is incredibly high exhaust velocity (though sometimes it can be too high).

    The disadvantage it the absurdly small thrust.

    To calculate the exhaust velocity of a Pure Fusion Rocket:

    Ve = sqrt( (2 * E) / m )


    • Ve = exhaust velocity (m/s)
    • E = energy (j)
    • m = mass of fuel (kg)

    Remember Einstein's famous e = mc2? For our thermal calculations, we will use the percentage of the fuel mass that is transformed into energy for E. This will make m into 1, and turn the equation into:

    Ve = sqrt(2 * Ep)


    • Ep = fraction of fuel that is transformed into energy
    • Ve = exhaust velocity (percentage of the speed of light)

    Multiply Ve 299,792,458 to convert it into meters per second.


    D-T fusion has a starting mass of 5.029053 and a mass defect of 0.018882. Divide 0.018882 by 5.029053 to get Ep of 0.00375.

    Plugging that into our equation Ve = sqrt(2 * 0.00375) = 0.0866 = 8.7% c. In meters per second 0.0866 * 299,792,458 = 25,962,027 m/s.

    Afterburner Fusion Engines

    Fusion afterburners use just the plasma thermal energy, but adds extra cold reaction mass to the fusion products.

    This is based on information from physicist Luke Campbell.

    For a given mission with a given delta V requirement, it is possible to calculate the optimum exhaust velocity. In many cases a fusion engine has thrust too low to be practical, but the exhaust velocity is way above optimal. It is possible to increase the thrust at the expense of the exhaust velocity (and vice versa) by shifting gears. An afterburner for a fusion engine is a way to shift gears.

    A pure fusion engine just uses the hot spent fusion products as the reaction mass. An afterburner fusion engine has a second plasma chamber (the afterburner) constantly filled with some cold propellant (generally hydrogen or water, but you can use anything that the spend fusion plasma can vaporize). The hot spent fusion products are vented into the afterburner, heating up the cold propellant. The average temperature goes down (decreasing the exhaust velocity) while the propellant mass flow goes up (increasing the thrust). The propellant mass flow increases naturally because instead of just sending the fusion products out the exhaust nozzle, you are sending out the fusion products plus the cold propellant. The contents of the afterburner are sent out the exhaust nozzle and Newton's Third Law creates thrust.

    In the equations below, a nozzle with an efficiency of 100% would have a efficiency factor of 2.0. But in practice the efficiency maxes out at about 85%, which has an efficiency factor of 1.7

    eq.1     Ptherm = F2 / (1.7 * (F / Ve))

    eq.2     mDot = F2 / (1.7 * Ptherm)

    eq.3     Ptherm = F2 / (1.7 * mDot)

    eq.4     F = sqrt[ 1.7 * Ptherm * mDot ]

    eq.5     Ve = F / mDot

    eq.6     mDot = F / Ve


    F = thrust (newtons)
    Ptherm = Thermal power (watts)
    mDot = propellant mass flow (kg/s) spent fusion product propellant + cold reaction mass
    Ve = Exhaust Velocity (m/s)
    1.7 = efficiency factor
    sqrt[ x ] = square root of x

    The thermal power is obtained from the fusion fuel table, using the % Thermal value. For instance, if you were using D + T fuel, 21% of the power from the burning fuel is what you use for Ptherm. That is, if the engine is burning 0.001 kilograms of D+T per second, it is outputting 339.72×1012 * 1×10-3 = 339.72×109 watts of energy, so Ptherm equals 339.72×109 * 0.21 = 7.1341×1010 watts.

    The amount of mDot contributed by spent fusion products can also be obtained from the fusion fuel table by using the TJ/kg column. For instance, with D+T fusion, if the rocket needs Ptherm of 2 terawatts, the total energy needed is 2 / 0.21 = 9.52 terawatts. The spent fusion products mDot is 9.52 / 339.72 = 0.028 kg/s. Usually the spent fusion product mass will be miniscule compared to the cold propellant mass. That is the reason the thrust was so miserably low to start with.

    The equation you use depends upon which value you are trying to figure out.

    1. When you have decided on the thrust and exhaust velocity, and want to know how much Thermal Power you need.
    2. When you have decided on the thrust and thermal power, and want to know how much propellant mass flow you need.
    3. When you have decided on the thrust and propellant mass flow, and want to know how much Thermal Power you need.
    4. When you have decided on the thermal power and the propellant mass flow, and want to know how much thrust you will get.
    5. When you have decided on the thrust and propellant mass flow, and want to know how much exhaust velocity you will get.
    6. When you have decided on the thrust and exhaust velocity, and want to know how much propellant mass flow you will need.

    Dual-Mode Fusion Engines

    Dual-mode use the neutron and bremsstrahlung radiation energy (which is otherwise wasted) to heat cold reaction mass, in parallel to the fusion products exhaust. In addition a Dual-mode can switch into Pure Fusion mode.

    This is based on information from physicist Luke Campbell.

    The neutron and bremsstrahlung energy produced by the fusion reaction is basically wasted energy when it comes to rocket propulsion. A dual-mode engine can switch from pure fusion mode into harvesting mode. This means additional cold propellant mass is moved around the fusion reaction chamber to be heated by the neutrons and bremsstrahlung radiation. This augments the thrust, at the expense of increasing the propellant usage rate.

    If the additional exhaust nozzles have an efficiency of 70%, and the additional propellant has an exhaust velocity of 10,000 m/s, the harvesting mode engine will create thrust of 1 newton per 7,000 watts of neutron + bremsstrahlung power, and consume 0.0001 kilograms of propellant per newton of thrust per second.

    There are some designs that try to harvest the wasted neutron and bremsstrahlung energy by attempting to turn it into electricity instead of thrust. But sometimes it is not worth it. To avoid excessive radiators the power generator typically have a maximum efficiency of 25% or less. So a maximum of 25% of the combined neutron+bremsstrahlung energy can be turned into electricity. This requires a turbine and electrical generator, which cuts into the payload mass.

    Breeder Fusion Engine

    This is from Maximizing Specific Energy By Breeding Deuterium (2019)

    Deuterium-Deuterium fusion is sort of the red-headed stepchild of fusion reactions. All the other reactions are easier to ignite (with the exception of hydrogen-boron) and produce more energy.

    In the following, "MeV" stand for Mega Electron-Volt. 1 MeV equals 1,000,000 electron-volts. Nuclear physicists use MeV a lot because it is a convenient size.

    Back in the 1970's researchers were taking a second look at D-D fusion. When you fuse two deuterons together, there are two fusion chains it can follow. One produces energy (4.0 MeV) plus one tritium nuclei and one hydrogen nuclei. The second chain produces energy (3.3 MeV) plus one helium-3 nuclei and one neutron. Hey, waitaminute! If you have some spare deuterium, the reaction products have the makings for the Deuterium-Tritium reaction and the Deuterium-Helium-3 reaction.

    So you fuse some deuterium, then send the reaction products into a second chamber with more deuterium and fuse that. Three deuterons are consumed. The end result is 21.6 MeV of energy; plus one Helium-4 nuclei (an alpha particle), one hydrogen nuclei, and one neutron. This is called the Catalyzed D-D fuel cycle and it is much better than the ordinary D-D reaction.

    In the paper, the author noted that the catalyzed reaction is still wasting energy on producing deadly neutrons. The neutrons produce zero thrust, and can damage the engine (and crew) three different ways. What if the hydrogen and neutrons (hydrogen from either the reaction products or extra carried on-board) were used to breed the hydrogen into more deuterium?

    All you have to do is surround the reaction chamber with a blanket of liquid hydrogen (inside a hollow-walled shell), let the neutron transmute hydrogen atoms into deuterium atoms, then skim out the deuterium (using the Girdler Sulfide process). The fresh deuterium can be sent back into the reaction chamber. The new fuel cycle is two deuterons are consumed. The end result is 23.8 MeV of energy, plus one Helium-4 nuclei. The paper calls this the Catalyzed D-D+D fuel cycle.

    Fusion Specific Energies
    Fuel CycleSpecific
    Catalyzed D-D+D
    breed reaction hydrogen
    6.0 MeV/AMUOpt 1: 11.3% c
    Opt 2: 4.4% c
    Opt 3: 7.2% c
    Catalyzed D-D+D
    breed carried hydrogen
    6.0 MeV/AMUOpt 4: 6.1% c
    Catalyzed D-D+D
    breed scooped hydrogen
    6.0 MeV/AMUOpt 5: 7.6% c
    Opt 6: 8.9% c
    D-3He3.7 MeV/AMU7.1% c
    Catalyzed D-D3.6 MeV/AMUOpt 0: 5.6% c
    D-T3.5 MeV/AMU
    D-T w/ T breeding2.8 MeV/AMU
    D-D0.9 MeV/AMU
    p-B-110.7 MeV/AMU

    The above table compares the specific energies of various fusion reaction cycles. That is, how much energy are they getting out of each atom? 1 AMU is more or less the mass of a hydrogen ion, MeV is how nuclear physicists like to measure nuclear energy.

    • p-B-11 is boron fusion, the only reason it's on the list is because the reaction produces no deadly neutrons. Specific energy wise it is the bottom of the list
    • D-D is the red-headed stepchild, barely above boron fusion in specific energy
    • D-T is next best specific energy. The only reason D-T w/ T breeding is on the list is because tritium rapidly decays in the fuel tanks and has to be restored
    • Catalyzed D-D makes the former red-headed stepchild even better than D-T in specific energy
    • D-3He was specific energy top-dog of the reasonable fusion engine fuels, the main problem is it's hard to manufacture and as far as we know is only found naturally in worthwhile amounts within gas giant atmospheres.
    • Catalyzed D-D+D turns Catalyzed D-D into the new specific energy top dog.

    The report is quick to note that the specific energy of a reaction is only one of many criteria to be considered when designing an engine.

    Engine Designs

    The report notes that "specific energy" is the mass of the fuel divided by the energy released. "Specific momentum" is the momentum of the exhausted particles divided by the mass of the input particles, so it is close to what we call the "exhaust velocity".

    The point is that a high specific momentum is important for pure fusion engines, while a high specific energy is important for afterburner fusion engines.

    Without deuterium breeding, the D-3He looks like it has a great specific energy and great specific momentum. And since all the reaction products are charged particles, they can be efficiently redirected to provide thrust. With Catalyzed D-D+D things are not so clear-cut. Some analysis is needed.

    OPTION 0: First they looked at the old-fashioned Catalyzed D-D fuel cycle, the one without the "+D" and no deuterium breeding. The full fuel cycle is:

    6 2H ⇒ 4He (3.5 MeV) + 1H (3.02 MeV) + 1n (14.1 MeV) + 4He (3.6 MeV) + 1H (14.7 MeV) + 1n (2.45 MeV) + 1.83 MeV

    • 1H = hydrogen
    • 2H = deuterium
    • 3H = tritium
    • 3He = helium-3
    • 4He = ordinary helium
    • 1n = neutron

    The kinetic energy of each particle is in parenthesis, and the energy carred by intermediate products which remain on-board are in the last term (1.83 MeV).

    This shows that fully 38% of the total energy is wastfully stolen by neutrons and cannot be used for thrust. Well, actually you can place a hemispherical neutron absorber shield ahead of the fusion reaction chamber. This will allow about 1/4 of the neutron energy to be used as thrust. This means only 28.5% of the total energy is wasted by neutrons.

    In addition, the energy in the tritium and helium-3 ions is wasted, because they have to be halted and sent to the secondary reaction chamber for further fusion. That wastes another 4% of the total energy.

    Bottom line is that Catalyzed D-D fuel cycle has a specific momentum of 5.6% c.

    OPTION 1: Adding the full coverage hydrogen blanket converts this into the Catalyzed D-D+D fuel cycle, the one with the "+D" and with deuterium breeding. The theoretical maximum is a specific momentum of 11.3% c. In practice it will be almost impossible to take all the 23.8 MeV of reaction energy and transfer it into accelerating the helium-4 ions. Especially since so much of the energy is in the form of rapidly moving neutrons.

    OPTION 2: With reaction hydrogen retained in order to breed more deuterium, the fuel cycle is:

    4 2H ⇒ 4He (3.5 MeV) + 4He (3.6 MeV) + 40.5 MeV

    Even assuming that the helium-4 is exhausted perfectally, the specific momentum is only 4.4% c. This is because the hydrogen produced in the fusion reaction has lots of momentum, all of which is lost when the hydrogen is halted and fed into the breeding blanket.

    OPTION 3: The fact remains that all Catalyzed D-D+D fuel cycles have 6.0 MeV/AMU, which is a heck of a lot. Compared to the others, this is tremendous amount of energy. Yes, braking to a halt the hydrogen loses about 20 MeV of momentum, but the braking process can convert the momentum into electricity. This can be used in an ion acceleration stage at the exhaust nozzle to add more speed to the helium-4 ions.

    IF the efficiency of transfering the 18 MeV to the helium-4 had an overall efficiency of just 30%, the 20 MeV would become about 6 MeV of acceleration. That would boost the helium-4 ions from 3.5 MeV to 9.5 MeV, resulting in a specific momentum of 7.2% c.

    OPTION 4: If you are not going to use an ion acceleration stage, the better option is to not halt the hydrogen reaction product. Let it contribute to the thrust in the exhaust. The drawback is the spacecraft will have to carry a tank of extra hydrogen for the breeding blanket.

    With reaction hydrogen jettisoned in exhaust, and extra hydrogen carried to breed deuterium:

    4 2H + 21H ⇒ 4He (3.5 MeV) + 1H (3.02 MeV) + 4He (3.6 MeV) + 1H (14.7 MeV) + 22.8 MeV

    This will give a specific momentum of 6.1% c. Better than standard Catalyzed D-D, but not as good as D-3He.

    OPTION 5: For some extreme speculation, use Option 4 but instead of carrying an onboard tank of hydrogen, give it a Bussard Ramscoop. That will let it scoop hydrogen out of the interstellar medium. This will raise the specific momentum from 6.1% to 7.6% c.

    The good news is that 22.8 MeV of on-board energy can energize a good-sized ramscoop. Which is a relief since such scoops are power hogs. Also Ramjets were thought to only be useful with impossible-to-do proton-proton fusion. But a Catalyzed D-D+D can make do with much easier D-D fusion.

    The bad news is that Bussards have some problems. Ramjets have a minimum speed to be effective at scooping, somewhere between 1% to 6% c. And the drag of the scoop will give the ship a terminal velocity equal to the exhaust velocity, which kind of negates the advantage of avoiding an on-baord hydrogen tank. On the other hand, this will work marvelously during the deceleration phase. There the drag will be to your advantage.

    OPTION 6: If you really want to go all-out, take the ramscooping Option 5, but take some of the energy away from the scoop and use it for an ion accelerator stage. Again with an overall efficiency of 30% you could boost the specific momentum up to 8.9% c.

    Nuclear Magnetic Spin Alignment

    This is an unobtainium way of turning a deuterium-tritium fusion reaction into a torch drive. You can find details here.

    ( STARFIRE Fusion Afterburner )

    This is a fictional fusion propulsion which is ingenious but probably impractical.

    Since the 1960s one of the leading schemes for controlling fusion, known as inertial confinement, had involved the implosion of tiny spheres of frozen hydrogen, spheres so small that hundreds could fit on the head of a pin—and every sphere a miniature H-bomb. The nation’s weapons laboratories, Los Alamos in New Mexico, Livermore in Califomia, Sandia in both states, had a monopoly on the classified knowledge essential to inertial confinement projects; who else regularly set off nuclear bombs and measured their behavior? Who else could generate mathematical models of nuclear explosions on the world's fastest computers?

    These diminutive superbombs were to be triggered by an array of powerful lasers or particle accelerators—ray guns, that is—arranged in a circle, pointing inward. Firing simultaneously, the beams would hit each frozen pellet as it fell into their midst. As the flash-heated surface expanded it would crush the sphere's interior until the hydrogen nuclei were fused into new elemental combinations—ideally releasing some three orders of magnitude more energy than that used to trigger the blast. Provided that it did not instantly melt the machine or blow it to pieces, this thousandfold increase in energy could be used to produce electrical power. Or to do other things. Fusion research had long been entangled with the military's yen for Buck Rogers-style death rays.

    That was okay with Linwood Deveraux. As reticent and gentle and genuinely polite as his soft Louisiana accent and his long-nosed, sad face suggested he was, Linwood nevertheless loved things that went zap and boom.

    His job, as one member of a brainy team at Livermore that called itself Q Branch, was to build an inertial confinement chamber that would convert thermonuclear explosions into directed beams of energy—ray guns thousands of times more powerful than those used as the spark plugs to ignite their hydrogen-pellet fuel.

    At the precise hour when the late afternoon photons came screaming through the westward window, bouncing off the neighbor’s asparagus fern in a blaze of light, tickling the shy lithops, punching him in the eye, Linwood got his modest idea.

    Several tricks were needed to design any fusion reactor, but they all required an intimate knowledge of the behavior of atoms and subatomic particles in the presence of strong electric and magnetic fields. That sort of knowledge, in turn, rested partly on a powerful intuition of geometry, and there is no useful theorem in geometry that cannot be at least qualitatively suggested with a paper and pencil. He sketched the lab’s current test machine,with its ring of lasers firing inward toward the tiny hydrogen pellet target and the strong magnetic “nozzle” that contained and directed the resulting explosion (Linwood Sketch 1).

    What is the heart of a thermonuclear explosion? Provided it starts and ends clean—uncontaminated by heavy elements like plutonium or uranium, which are intrinsic to the brute force of real H-bombs—a thermonuclear explosion is a clear hot soup, a plasma of protons, electrons, ionized helium, free neutrons, un-ionized hydrogen atoms and leftover neutrinos and such. All the electrically charged particles will stay in the soup, if it is confined and shaped by electric and magnetic fields.

    Most of the energy of the explosion, more than three quarters, is in the form of speeding neutrons. Neutrons aren’t significantly affected by electric or magnetic fields, but they can be slowed in a materially dense "blanket”—liquid lithium or some such substance—their energy thus converted to heat.

    What happens to the heat depends on what the reactor is designed to do. A power reactor uses it to make steam, and eventually electricity, but in ray guns most of the heat is a waste and a nuisance. Over the decades ray gun designers had played with various ways of using the energy of a thermonuclear explosion, for example, by letting it squeeze magnetic fields to produce huge electrostatic charges, or by opening one end of the magnetic bottle to let the products spew out, or by focusing some of the energy into x-ray beams, and so on—but no matter what the scheme, most of the thermonuclear reactor’s energy was wasted as heat.

    On his sketch pad, Linwood roughed in the liquid blanket and the circulating coolant systems required to dispose of the waste heat (Linwood Sketch 2). There he paused.

    To Linwood, with his passion for efficiency, wasting so much heat had always seemed criminal. Surely something clever could be done with those copious neutrons!

    In power reactors, neutrons were intrinsic to the fusion fuel cycle; they were captured to breed radioactive tritium, the rarer (because of its short half-life) of the two isotopes of hydrogen that composed the fusion fuel, the other being the more common deuterium. Tritium breeding was a secondary process, however, a civilian process. A ray gun orbiting in space would be supplied with all the tritium it was ever likely to need.

    Linwood wondered about other neutron-capture scenarios. In the heart of the sun, neutron capture contributed to the formation of heavier elements…but to take advantage of stellar fusion processes was a dream of the far future, awaiting the day when truly monstrous magnetic fields could be generated, capable of rivaling gravity at the heart of a star.

    Linwood thought about all this a long time and drew nothing. Applying the old creative principle that when the going gets tough the smart go somewhere else, he balled up his rough sketch and threw it away.

    He stared morosely at the lithops, now faintly glowing in the setting sun's last light.

    At the lab his discarded sketch would have been sucked into a high-temperature furnace and instantly reduced to fine ash, but at home Linwood had an open wastebasket beside his table, the contents of which he conscientiously burned…whenever he remembered to. The security disadvantages of this practice were offset by certain practicalities, one of which Linwood now demonstrated to himself—

    —by changing his mind. He fished the crumpled wad of paper out of the basket and flattened it on his table. Now what if, instead…

    The lithium-neutron reaction that yields tritium is quite efficient: a neutron entering a blanket of liquid lithium travels ten or twenty centimeters and scatters from a few lithium atoms, heating up the neighborhood before strongly interacting with one of them to create a helium nucleus and a tritium nucleus. A typical power reactor circulates the liquid metal lithium through heat exchangers, meanwhile tapping a small side flow from which the tritium is chemically extracted.

    Instead of thinking of the lithium blanket as a coolant, optional for his purposes, Linwood tried thinking of it as an extra fuel tank. He imagined introducing lithium into an annular ring around the reaction chamber at a steady rate, letting it circulate long enough to be bombarded by sufficient neutrons to produce a good proportion of tritium. He imagined mixing this tritium-enriched fluid with a separate supply of deuterium. He imagined injecting the lithium-tritium-deuterium mix into a magnetically confined and compressed outflow of hot plasma from the primary reactor—in such a manner that it burst into a secondary fusion reaction, additionally heating an outgoing beam (Linwood Sketch 3).

    Hot stuff. Not all that efficient in the long run—only a small fraction of the injected fuel would fuse, even under ideal circumstances, and a great deal of waste heat would still have to be disposed of by radiators—but Linwood was satisfied that at least he had salvaged some neutrons.

    It was dark outside when Linwood happily finished his sketching. Not that he made an improvement on the Q Branch beam projector; he knew that what he'd drawn had little or nothing to do with death rays. That was fine with him. He turned out the light and went upstairs, made himself a wiener sandwich, and lay down in bed, after popping a chip into the viddie—that classic British thriller from the 1950s, X the Unknown, starring Dean Jagger; it had a great monster, a puddle of smart, ravenous, radioactive mud.

    The next day, when Linwood displayed a suitably gussied—up draft of his idea to his coworkers in Q Branch, the youngest of them, a kid on summer loan from MIT, had an attack of giggles. What Linwood had drawn had nothing to do with ray guns, said the pimpled kid—who spoke as an authority on ray guns, having read in his short life a great deal of space opera and very little else—and it wasn't even all that original. What ol' Linwood had here was an afterburner for a fusion rocketship.

    From STARFIRE by Paul Preuss (1988)

    ( AV:T Fusion )

    AV:T Fusion
    Cruise mode
    Exhaust Velocity832,928 m/s
    Specific Impulse84,906 s
    Thrust245,250 N
    Thrust Power0.1 TW
    Mass Flow0.29 kg/s
    Combat mode
    Exhaust Velocity104,116 m/s
    Specific Impulse10,613 s
    Thrust48,828,125 N
    Thrust Power2.5 TW
    Mass Flow469 kg/s

    Fictional magnetic bottle fusion drive from the Attack Vector: Tactical wargame. It uses an as yet undiscovered principle to direct the heat from the fusion reaction out the exhaust instead of vaporizing the reaction chamber. Like the VASIMR it has "gears", a combat mode and a cruise mode. The latter increases specific impulse (exhaust velocity) at the expense of thrust.

    In the illustration, the spikes are solid-state graphite heat radiators, the cage the spikes emerge from is the magnetic bottle, the sphere is the crew quarters and the yellow rectangles are the retractable power reactor heat radiators. The ship in the lower left corner is signaling its surrender by deploying its radiators.

    ( THS Fusion Pulse )

    Fusion Pulse low gear
    Exhaust Velocity150,000 m/s
    Specific Impulse15,291 s
    Thrust80,000 N
    Mass Flow0.53 kg/s
    Fusion Pulse high gear
    Exhaust Velocity300,000 m/s
    Specific Impulse30,581 s
    Thrust40,000 N
    Mass Flow0.13 kg/s
    Thrust Power6.0 GW
    Total Engine Mass4,000 kg
    Specific Power1 kg/MW

    Fictional inertial-confinement fusion drive from the game GURPS: Transhuman Space. Like the VASIMR it has "gears", one increases specific impulse (exhaust velocity) at the expense of thrust.

    ( Epstein Drive )

    Epstein Drive
    Thrust Power5.5 TW
    Exhaust Velocity11,000,000 m/s
    Specific Impulse1,100,000 s
    Thrust1,000,000 N
    Mass Flow0.09 kg/s

    Fictional Magnetic Confinement Fusion drive from The Expanse series. The sparse details I managed to find were from the short story Drive.

    The inventor mounted the newly-invented drive in a small interplanetary yacht whose living space was smaller that Epstein's first Mars apartment. When the fuel/propellant tanks were 90% full, the drive could produce 68 m/s2 acceleration (6.9 g). Which was quite a few times higher than Epstein was expecting. He was instantly pinned by the acceleration and could not turn the drive off. The drive burned until the tanks were dry, which took 37 hours and had delta-V'd the yacht up to 5% c (roughly 15,000,000 m/s). By this time Epstein was long dead and the yacht can still be seen by a powerful enough telescope on its way to nowhere.

    The drive was some species of fusion drive using Epstein's innovative "magnetic coil exhaust". The yacht started with propellant tanks 90% full. After 10 minutes they had dropped to 89.6% full. After 2 more minutes 89.5%. After 2.5 more minutes 89.4%. After 37 hours 0% full.

    Thus ends the canon knowledge.

    My Analysis

    Now comes conjecture on my part. Please note this is totally non-canon and unofficial, I'm just playing with numbers here.

    I made lots of assumptions. I assumed the yacht had a mass ratio of 4, since Jerry Pournelle was of the opinion that was about the maximum for an economical spacecraft. I also assumed the yacht had a mass of 15 metric tons, because that was the wet mass of the Apollo Command and Service module.

    What does those assumptions give us?

    If the delta V is 5% c and the mass ratio is 4, the exhaust velocity has to be about 11,000,000 m/s, or 3.7% c. ( Ve = ΔV / ln[R] )

    Looking over the theoretical maximum exhaust of various fusion reactions we find we are in luck. Pretty much all of them can manage more than that exhaust velocity, with the exception of Deuterium-Helium3.

    Given an acceleration of 68 m/s2 and estimated wet mass of 15,000 kg, the thrust has to be 1,000,000 Newtons. ( F = Mc * A ). For one engine.

    If we use the estimated thrust of 1,000,000 Newtons and estimated exhaust velocity of 11,000,000 m/s, the propellant mass flow is an economical 0.09 kg/s. ( mDot = F / Ve )

    Of course the thrust power is a whopping 5.5 terawatts, but what did you expect from a torchship? ( Fp = (F * Ve ) / 2 )

    Feel free to make your own assumptions and see what results you get.

    Scott Manley's Analysis

    The legendary Scott Manley does his own analysis of Epstein's experimental ship in this video. He figures that: Yes a fusion drive will give the needed performance but No the heat from the drive will vaporize the entire ship in a fraction of a second.

    Monstah's Analysis

    Independently of assuming a specific ship's mass and propellant fraction, he takes the hard canon facts of Epstein's experimental ship having an acceleration of 6.9 gees and a delta V of 5% c, and calculates a result of an exhaust velocity of 13,000,000 meters per second and a mass ratio of 3.0 to 3.3.

    Start with mass ratio equation

    R = M / Me = (Mpt + Me) / Me

    where Me and Mpt are dry and propellant masses.

    Now, substitute an expression for propellant mass

    Mpt = mDot * t

    where mDot is the mass flow and t the time till total consumption (t=37 hours is given in the problem).

    Mass flow mDot can be calculated from thrust and exhaust velocity

    mDot = F / Ve

    Thrust (and fuel flow) can be assumed constant; calculated at the initial time, it's

    F = m * A

    for A = 68 m/s2 (6.9 gees) and m the initial mass (same used for mass fraction, M)

    We now have

    R = (Me + Mpt) / Me = (Me + (mDot * t)) / Me = (Me + (m * A / Ve) * t) / Me

    The equation above simplifies to

    1 / (1-(t*A / Ve)) = R = expV / Ve)

    where ΔV is 5% c given

    We now have an equation with a single variable, Ve! However, it's an ugly ass equation where Ve appears both as a denominator in an exponent and a denominator in a nested fraction. Ew.

    Wolfram Alpha to the rescue! \o/

    Telling it to solve for Ve, we get

    Ve = A * ΔV * t / (A * t * productlog(-ΔV / (A*t) * exp(-ΔV / (A*t))) + ΔV)

    where productlog() is the "ProductLog function". Don't ask.

    If you just plug in the values where they appear you'll get a timeout, so I'll precalculate A*ΔV * t, A*t and ΔV / (A*t), convert everything to meters and seconds and ignore the units, and throw in WolframAlpha again.

    The answer is Ve ~ 13,000 km/s (13,000,000 m/s or 4.3% c).

    Very close to your 11,000 km/s (but, importantly, independently of any assumptions of ship mass and fuel fraction). You assumed R = 4, the result here is closer to 3. But then, our initial time had the ship at 90% propellant tank capacity, so the ship's actual design is for something around Mass Ratio 3.3

    From MONSTAH

    Erin Schmidt's Analysis

    Erin Schmidt did a quick analysis of the Epstein-drive ship Rocinate (not Epstein's experimental ship), hinging on some very loose assumptions. He figures the thrust power is 11 terawatts. Egads.

    Mass Ratio R = 3.0
    Dry Mass Me = 500,000 kg

    NOVEL STATES Rocinante can accelerate at 0.25 g for 3 to 4 weeks (2.45 m/s2 for 2.419×106 seconds)
    2.45 * 2.419×106 = 5,933,000 m/s = 6000 m/s delta-V

    Specific Impulse
    Isp = (ΔV / ln(R)) / g0
    Isp = (5,933,000 / ln(3.0)) / 9.81
    Isp = 551,000 seconds

    Exhaust Velocity
    Ve = ΔV / ln(R)
    Ve = 5,933,000 / 1.0986
    Ve = 5,400,000 m/s = 0.018c = 18% c

    Wet Mass
    M = R * Me
    M = 3.0 * 500,000
    M = 1,500,000 kg

    F = M * 0.25 * g0
    F = 1,500,000 * 0.25 * 9.81
    F = 3,680,000 Newtons = 3700 kN

    Thrust Power
    Fp = (F * Ve ) / 2
    Fp = (3,700,000 * 6,000,000 ) / 2
    Fp = 11,100,000,000,000 Watts = 11 TW

    Matter Beam's Analysis


          We aim to take a fictional propulsion technology from The Expanse, and apply the appropriate science to explain its features in a realistic manner.

         This also applies to other SciFi settings that want a similar engine for their own spacecraft. The Epstein Drive

         Title art is from here.

         Central to the setting of The Expanse is a very powerful fusion-powered engine that allows spacecraft to rocket from one end of the Solar System to the other quickly and cheaply.

         It reduces interplanetary trips to days or weeks, allows small shuttles to land and take off from large planets multiple times and accelerate at multiple g’s for extended amounts of time.

         Such a propulsion system is known as a ‘torch drive’: huge thrust, incredible exhaust velocity and immense power inside a small package.

         Fusion energy can certainly provide these capabilities. Using fuels like Deuterium provides over 90 TeraJoules of energy per kilogram consumed. Proton fusion, the sort which powers our Sun, could release 644 TJ/kg if we could ever get it to work.

         The Epstein Drive (art Gautam Singh) is described in The Expanse as a breakthrough in fusion propulsion technology. A short story provides some details. A small spaceship equipped with this engine could reach 5% of the speed of light in 37 hours, averaging over 11g’s of acceleration. A magnetic bottle is mentioned. Since we don’t know the mass of the vehicle or what percentage of it was propellant, we can’t work many useful details.

         The main book series and the TV show focus on the adventures of the Rocinante and its crew. We know that it uses laser-ignited fusion reactions and water as propellant. Again, we don’t have a mass or propellant fraction, so we cannot get definitive performance figures. However, we have detailed images of its interior and exterior. Note that there are no radiator fins or any heat management system visible.

         The cross-section also reveals that there is very little room for propellant. Despite this, it can accelerate at over 12g’s and has reached velocities of 1800km/s while averaging 5g. Presuming that it can slow down again and jet off to another destination, this implies a total deltaV on the order of 4000km/s, which is 1.33% of the speed of light.

         Official figures for the masses of spacecraft from The Expanse do exist. In collaboration with the TV show’s production team, SpaceDock created a series of videos featuring ships such as the Donnager-class Battleship for which a mass of 250,000 tons is provided.

         Using the battleship’s dimensions, we obtain an average density of about 20 to 40 kg/m^3. For comparison, the ISS has a density of 458 kg/m^3. We will use this average density for now, but you can read the Scaling section below to understand how different mass assumptions for the Rocinante don't 'break' our workings so far. 

         Applying the battleship density to the Rocinante's size gives us a mass of about 130 to 260 tons. It is likely to change a lot depending on what the ship is loaded with, seeing as it is almost entirely made up of empty volumes. We’ll use a 250 tons figure for an empty Rocinante and add propellant to it as needed.

         Let’s put all these numbers together.

         The Epstein drive technology allows for >250 ton spacecraft to accelerate for several hours at 5g with bursts of up to 12g, achieving a deltaV of 4000km/s, while not having any radiators and a tiny propellant fraction.

         Can we design a realistic engine that can meet these requirements?

    The Heat Problem

         The biggest problem we face is heat.

         No engine is perfectly efficient. They generate waste heat. Some sources of waste heat are physically unavoidable, however performant the machinery becomes.

         Fusion reactions result in three types of energy: charged kinetic, neutron kinetic and electromagnetic.

         Charged kinetic energy is the energy of the charged particles released from a fusion reaction. For a proton-Boron reaction, it is the energy of the charged Helium ions (alpha particles) that come zipping out at 4.5% of the speed of light.

         We want as much of the fusion reaction to end up in this form. Charged particles can be redirected out of a nozzle with magnetic fields, which produces thrust, or slowed down in a magnetohydrodynamic generator to produce electricity. With superconducting magnets, the process of handling and using charged kinetic energy can be made extremely efficient and generate practically no heat.

         Neutron kinetic energy is undesirable. It comes in the form of neutrons. For deuterium-tritium fusion, this represents 80% of the fusion output. We cannot handle these particles remotely as they have no charge, so we must use physical means. Neutron shields are the solution; the downside is that by absorbing neutrons they convert their all of their energy into heat. This is a problem because materials have maximum temperatures and we cannot really use radiator fins to remove the heat being absorbed.

         Electromagnetic radiation is another unavoidable source of heat. Mirrors can reflect a lot of infrared, visual and even ultraviolet wavelengths. However, fusion reactions happen at such a high temperature that the majority of the electromagnetic radiation is in the form of X-rays. These very short wavelengths cannot be reflected by any material, and so they must also be absorbed.

         With this information, we can add the following requirements:

    • We must maximize energy being released as charged particles.
    • We must minimize heat from neutron kinetic and electromagnetic energy.

         Thankfully, there is a fusion reaction that meets these requirements.

    Diagram from here

         Helium-3 and Deuterium react to form charged Helium-4 and proton particles. Some neutrons are released by Deuterium-Deuterium side reactions, but by optimizing the reaction temperature, this can be reduced to 4% of the total output. An excess of Helium-3 compared to Deuterium helps reduce the portion of energy wasted as neutrons down to 1%. Another 16% of the fusion energy becomes X-rays. Other ‘cleaner’ source of fusion energy exists, using fuels such as Boron, but they cannot be ignited using a laser.

         An optimized Deuterium and Helium-3 reaction therefore releases 1 Watt of undesirable energy (which becomes waste heat if absorbed) for every 4 Watts of useful energy.

         If this reaction takes place inside a spaceship, then all of the undesirable energy must be turned into heat. However, if it is done outside the spaceship, then we can get away with only absorbing a fraction of them. It's the idea behind nuclear pulsed propulsion. 

         How else do we reduce the potential heat a spaceship has to absorb?


         A fusion reaction produces a sphere of very hot plasma emitting neutrons and X-rays in all directions. A spaceship sitting near the reaction would eclipse most of these directions and end up absorbing up to half of all this undesirable output.

         If the fusion reaction takes place further away, less of the undesirable output reaches the spaceship and more of it escapes into space.

         It is therefore a good design choice to place the fusion reaction as far away as possible. However, we are limited by magnetic field strength.

         The useful portion of the fusion output, which is the kinetic energy of the charged particles, is handled by magnetic fields to turn it into thrust. Magnetic fields quickly lose strength with distance. In fact, any magnetic field is 8 times weaker if distance is merely doubled. 10 times further away means a field a 1000 times weaker. If we place the fusion reaction too far away from the source of these magnetic fields, then the useful fusion products cannot be converted into thrust.

         We could calculate exactly how far the fusion reaction could take place from the spaceship while still being handled by magnetic fields, but whether you use magnetic beta (magnetic pressure vs plasma pressure) or the ion gyroradius (turning radius for fusion products inside a magnetic field), it is clear that kilometres are possible with less than 1 Tesla. For a setting with the Expanse’s implied technology level, generating such field strengths is easy.

         What does this all mean for a fusion engine?

         If we can generate a magnetic field strong enough to deflect fusion particles at a considerable distance, then we can convert a large fraction of the fusion output into thrust while only a small fraction of harmful energies reaches the spaceship.

         The Rocinante is about 12 meters wide. If we describe it as a square, it has a cross-section of about 144m^2. A fusion reaction taking place 20 meters away from the spaceship would have spread its undesirable energies (neutrons and X-rays) over a spherical surface area of 5027m^2 by the time they reach the Rocinante. This means that 144/5027= 2.86% of the fusion reaction’s energy is actually intercepted by the spaceship.

         Increase this distance to 200 meters and now only 0.0286% of the fusion reaction’s harmful output reaches the spaceship. A much more powerful fusion output is possible.

         Finally, we need a heatshield.

    NASA heatshield materials test.

         Despite only a portion of the fusion output being released as neutrons and X-rays, and a small fraction of even that becoming radiation that actually reaches the spaceship, it can be enough to melt the ship.

         We therefore need a final barrier between the fusion reaction and the rest of the spaceship. A heatshield is the solution.

         This heatshield needs to enter into a state where it balances incoming and outgoing energy. With no active cooling available, no heatsinks or external fins, the heatshield has to become its own radiator.

         The Stefan-Boltzmann law says that a surface can reach the state described above at its equilibrium temperature. It can be assumed that emissivity is high enough to not matter (over 0.9).

         Equilibrium temperature = (Incoming heat intensity/ (5.67e-8))^0.25

    Equilibrium temperature is in Kelvin.
    The heat intensity is in Watts per square meter (or W/m^2)

         Using this equation, we can work out that an object sitting under direct sunlight in space (at 1 AU from the Sun, so receiving 1361 W/m^2) would have an equilibrium temperature of 393 Kelvin.

         A concentrating mirror focusing sunlight to 1000x intensity (to 1.36 MW/m^2) would heat up an object to the point where radiates heat away at a temperature of 2213K.

         For a fusion-powered spaceship, you want this temperature to be as high as possible. Higher temperatures means that the incoming heat intensity can be greater, which in turn means that the spaceship can shield itself from more powerful fusion outputs.

         Tungsten, for example, can happily reach a temperature of 3200K and survive a beating from 5.95MW/m^2.

         Graphite can handle 3800K before it starts being eroded very quickly. That’s equivalent to 11.8MW/m^2.

         Tantalum Hafnium Carbide is the current record holder at 4150K. Keep it below its melting temperature at 4000K, and we would see it absorb 14.5MW/m^2. Scientists have also simulated materials which could reach over 4400K before they melt.

         This heatshield needs to rest on good insulation so that it doesn’t conduct heat into the spaceship. A design similar to the Parker Solar Probe’s heatshield mounting can be used. Low thermal conductivity mountings and low emissivity foil can reduce heat transfer to a trickle.

    Proposed design

         Let’s talk specifics.

         We will describe now a fusion-powered rocket engine design that can perform most similarly to the Expanse’s Epstein Drive as shown on the Rocinante.


         It is based on this refinement to the VISTA fusion propulsion design. Like the VISTA design, a laser is used to ignite a fusion fuel pellet at a certain distance from the ship and a magnetic coil redirects the fusion products into thrust. The rear face of the spaceship takes the full brunt of the unwanted energies and re-emits them as blackbody thermal radiation.

         The refinement consists of a shaped fusion charge that can be ignited by laser slamming a portion of the fusion fuel at high velocity into a collapsing sphere, raising temperatures and pressures up to ignition levels.

         Instead of the fusion products being released in all directions, a jet of plasma is directed straight at the spaceship. This increases thrust efficiency up to 75%, as the paper cites.

    Somewhat similar magnetic nozzle configuration from this MICF design

         Meanwhile, the X-rays and neutrons escape the plasma in all directions.

         The Epstein Drive is assumed to be a version of this. Instead of a spherical firing squad of lasers (as can be found in the NIF facility) that requires lasers to be redirected sideways with mirrors, a single laser is used for ignition. It is less effective but it means we can dispense with mirrors hanging in space. 

         We will also be using Deuterium and Helium-3 fuels instead of Deuterium and Tritium. They are harder to ignite, but give much more useful energy (79% comes out as charged particles). By adjusting the fusion temperature and ratio of Helium-3 to Deuterium, we can increase this output to become 83% useful while neutrons fall to 1% of the output and X-rays represent 16%.

         Also, using powerful magnetic coils, we will be igniting the fusion pellets at a much greater distance from the physical structures of the engine. We can take the 'nozzle' to actually be a mounting for the magnetic coil and everything with a line of sight to the fusion reaction to be covered in a heatshield. More importantly, the engine will be much, much smaller than the 120m diameter of the VISTA design.

         The Rocinante has a cross-section area of 144m^2. Its heatshield will be a black metal carbide that can reach an equilbirium temperature of 4000K. It is separated from the hull with insulating brackets that massively reduce the heat being conducted to the 300K interior.

         The heatshield needs to be thick enough to fully absorb X-rays and neutrons from the fusion reaction (it might be supplemented by boron carbide in cooler <3000K sections).

         At 4000K, the heatshield can handle 14.5MW/m^2. The rear of the Rocinante can therefore absorb 2.09GW of heat.

         The magnetic field acts on a fusion reaction 300m away from the hull. It acts like a spring; it requires no energy input to absorb the kinetic energy of the charged fusion reaction products and transmit it to the spaceship. Using figures from the cited paper, thrust efficiency is 75%.

         Thanks to this arrangement, only 0.0127% of the unwanted energies from the fusion reaction are intercepted and absorbed as waste heat by the heatshield.

         Some of the heat can be converted into electricity and used to power the laser igniting the fusion reactions. The generator can be of the superconducting magnetohydrodynamic type, and the laser could be cryogenically cooled. This makes them both extremely efficient. The electrical power that needs to be generated to run the laser can be very small if the fusion gain is extreme (small ignition, big fusion output).

         Putting these percentages so far together, 0.00216% of the fusion reaction energy ends up as heat in the heatshield.

         Using that percentage, it is now evident that we have a very large ‘multiplier’ to play with. For every watt that the heatshield can survive, 46,300 watts of fusion output can be produced.

         A heatshield absorbing 2.09 GW of heat means that its Epstein drive can have an output of 96.8 TW. About 2.2kg of fuel is consumed per second.

         83% of that fusion power is in the form of useful charged particles, and the magnetic field turns 75% of those into thrust. So, 41.5% of the fusion power becomes thrust power; which is 60.25 TW.

         The effective exhaust velocity of a Deuterium-Helium3 reaction can be as high as 8.9% of the speed of light. This assumes 100% burnup of the fusion fuel. Because we are using an excess of Helium-3, this might be reduced to 6.3% of the speed of light.

         With this exhaust velocity, we get a thrust of 6.37 MegaNewtons.

         An empty 250 ton Rocinante would accelerate at 2.6g with this thrust.

         We know it can accelerate harder than that but it cannot handle any more fusion power. So, it must increase its thrust by injecting water alongside fuel into its exhaust.

         There is a linear relationship between exhaust velocity and thrust at the same power level, but a square relationship between thrust and mass flow.

         Halving the exhaust velocity doubles the thrust but quadruples the mass flow rate. The Rocinante can have a ‘cruise’ mode where only fuel is consumed to maximize exhaust velocity, and a ‘boost’ mode where more and more water can be added to the exhaust to increase thrust.

         It is useful to know this, as we must now work out just how much fuel (Deuterium and Helium-3) and extra propellant (water) it needs.

         1800km/s is done in the ‘boost’ mode, and then 2200km/s in the ‘cruise’ mode, for a total of 4000km/s. How much fuel and propellant does it need?

         As with any rocket equation calculation, we need to work backwards.

         Mass ratio = e^(DeltaV/Exhaust Velocity)

         An exhaust velocity of 6.3% of the speed of light and a deltaV requirement of 2200km/s means a mass ratio of 1.123.

         The 250 ton Rocinante needs to first be filled with 30.75 tons of fusion fuel. A 1:2 mix of Deuterium (205kg/m^3) and Helium-3 (59kg/m^3; it won't freeze) has an average density of 107.6kg/m^3, so this amount of fuel occupies 285m^3. It represents about 4.9% of the spaceship’s 12x12x40 m internal volume.  

         And now the ‘boost’ mode. 5g of acceleration while the spaceships gets lighter as propellant is being expended means that thrust decreases and exhaust velocity increases gradually over the course of the engine burn. The propellant load can to be solved iteratively... on a spreadsheet.

         Using 0.25 ton steps for water loaded onto the Rocinante, it can be worked out that an initial mass of 352 tons is required. This represents an additional 57 tons of water and 17.25 tons of fuel.

         The full load is therefore 57 tons of water in 57m^3, and 48 tons of fusion fuel in 446m^3. Together, they fill up 8.7% of the Rocinante’s internal volume.

         The thrust level during the acceleration to 1800km/s varies between 13.77MN and 17.27MN. It takes just over 10 hours to use up all the water.

         Boosting to 12g would require that this thrust be increased further, between 33.05MN and 41.44MN. However, it could only be sustained for 106 minutes, until 751km/s is reached.

    Official art by Ryan Dening

         In ‘cruise’ mode and with no water loaded, the Rocinante would have 3320km/s of deltaV and can cross the distance between Earth and Saturn in 10 to 12 days at any time of the year.

         At 12g, it can sprint out to a distance of 21.2 thousand kilometres in about 10 minutes, and 0.76 million kilometres in an hour.


         This proposed design can be easily scaled to adjust for different figures for mass, acceleration and deltaV.

         The variable will be the ignition point distance from the spaceship and therefore the magnetic field strength of the coils in the 'nozzle'. A stronger field allows for fusion products to be redirected from further away, so that an even smaller portion of the harmful energies are intercepted.

         If we assumed a ten times greater density for the Rocinante, for example, we would have an empty mass of 2500 tons. To adjust for this while maintaining the same performance, we would simply state that the fusion reaction is ignited 10^0.5: 3.16 times further away, or 948m. The 'multiplier' mentioned earlier jumps from 46,300 to 461,300, just over 10 times better than before. In other words, the fusion output can be increased 10 times and all the performance falls back in line with what was calculated so far.


         Beyond what we’ve seen on the show or read from the books, there could be some interesting consequences to having this sort of design.

         Visually, for example, the rear end of spaceships would glow white hot. They cannot come close to each other while under full power, as then they’d expose each other’s flanks to intense heat from the fusion reactions.

         You might have noticed from a previous diagram that a portion of the fusion plasma travels all the way up the magnetic fields without being redirected. This could be the reason why we see 'gas' in the 'nozzle'; it is simply the leaking plasma hitting a physical structure and being compression heated up to visible temperatures. 

         A failure of the magnetic fields would immediately subject the heatshield to 5x its expected heat intensity. This would quickly raise the temperature by a factor 2.23, so it would turn from solid to explosively expanding gas. Not exactly a ‘failure of the magnetic bottle’, but a similarly devastating result.

         On the other hand, the magnetic field passively provides shielding against most of the radiation that can affect space travellers. If it is strong enough to repel fusion protons, then it could easily deal with solar wind protons and other charged particles, as found in the radiation belts around Earth or Jupiter. This could be a reason why we don’t see thick blankets of radiation shielding all around the hull.

         Our proposed engine design is pulsed in nature. We want smooth acceleration, so we want as many small pulses in such quick succession that the spaceship feels a near-continuous push. This can be achieved with as few as 10 pulses per second, or hundreds if possible.

         However, even at 10 pulses per second, you need to shoot your fusion fuel from the fuel stores to the ignition point 300 meters away at a velocity of 3km/s. This can be accomplished by a railgun, and it is incidentally a good fraction of the projectile velocities used in combat.

         Could the Belters in the Expanse simple have pointed their fuel injection railguns in the opposite direction to equip themselves with their first weapons?

         Similarly, an intense laser is needed to ignite the fuel quickly enough to achieve an extreme fusion gain. Doing so from 300m away requires a short wavelength and a focusing mirror… which are also the components needed to weaponize a laser. If a laser can blast a fuel hard enough to cause it to ignite, then it could do the same to pieces of enemy spaceships, and all that is needed to extend the range is a bigger mirror. This implication can be countered by having an extreme fusion gain ratio — i.e., a 10,000,000 fold ratio between the energy input of a laser ignition system to the fusion output. That means a 100TW reaction can be ignited by just a 10MW laser, which is far less likely to be weaponized.

         There is also a claim made where the Rocinante’s fuel reserves are ‘enough for 30 years’. This cannot mean propulsion. Even at a paltry 0.1g of acceleration in ‘cruise’ mode, the Rocinante can consume all of its fuel in just 40 days. Add in a lot of drifting through space without acceleration, and we’re still looking at perhaps a year of propulsion. It is much more likely that this claim refers to running the spaceship; keeping the lights on, the life support running and the computers working. That sort of electrical demand is easily met by the energy content of fusion fuels.

         Finally, keep in mind that the propulsion technology described here is not specific to the setting of the Expanse. It respects physics and you can introduce it to any setting where real physics apply. In other words, it is a ready-made and scalable solution for having rapid travel around the Solar System without much worry about propellant, radiators, radiation shielding and other such problems!

    From THE EXPANSE'S EPSTEIN DRIVE by Matter Beam (2019)

    Nuclear Thermal

    Basically a Nuclear Thermal Rockets (NTR) is a nuclear reactor where the propellant is the coolant. And instead of the coolant being directed into a cooling tower, it instead exits out the exhaust nozzle, creating thrust.

    They use the heat generated from a nuclear reaction to heat up propellant. The nuclear reaction is controlled by adjusting the amount of free neutrons inside the mass of fissioning material (like all nuclear reactors do, generally with reactor control drums).

    As a side effect, if you have a cluster of several such engines it is vitally important to have distance and neutron isolation shields between them. Otherwise the nuclear reaction in each engine will flare out of control due to the neutron flux from its neighbor engines.

    The fact that the propellant is also the coolant means that after a thrusting period is over, you still have to vent propellant through the reactor after you turn it off. Until the reactor goes cold.

    Exhaust Velocity Limits on Nuclear Thermal Rockets

    The exhaust velocity and specific impulse of NTR are proportional to the thermal levels inside the reactor. Which a fancy way to say "the hotter the reactor, the higher the exhaust velocity."

    Which brings us to the exhaust velocity limit. Solid core NTRs use a solid-core nuclear reactor. Such reactors are made of matter. And as with all matter, if you raise the temperature, at some point it will get hot enough so that the reactor melts. Which means the core ain't solid any more. This is a bad thing, technical term is Nuclear meltdown, non-technical term is The China Syndrome. The molten remains of the reactor shoots out the exhaust bell like a radioactive bat from hell, killing anybody nearby and leaving the spacecraft without an engine.

    To avoid this unhappy state of affairs, solid core NTRs are limited to a temperature of about 2,750 K (4,490° F), which translates into an exhaust velocity limit of about 8,093 m/s (with liquid hydrogen, double that if you've manage to figure out how to stablize monoatomic hydrogen). Some fancy high temperature designs can push that up to an exhaust velocity of about 11,800 m/s.

    Lateral thinking rocket engineers had the brainstorm of "what if the reactor starts out molten in the first place?" This lead to the design of liquid-core NTR, with a temperature of 5,250 K and an exhaust velocity of 16,000 m/s.

    Because rocket engineers can't resist turning it up to 11, they figured if liquid is good then gaseous should be even better. This is the open-cycle gas-core NTR, with an exhaust velocity of a whopping 34,000 m/s.

    The major draw-back of open-cycle GCNTR is that there is no feasible to prevent any of the radioactive fission products and unburnt uranium from escaping out the exhaust. Which more or less makes the exhaust plume a weapon of mass destruction, and significatly increases the radiation exposure on the poor ship's crew.

    Engineers tried to fix the radiation problem of the open-cycle GCNTR by making it closed-cycle; that is, preventing physical contact between the gaseous uranium and the propellant. This turned out to be an attempt to have your cake and eat it too. The entire point of gas core was to allow outrageous engine temperatures by not having any solid components inside the engine, but sadly baffles that prevent the uranium from mixing with the propellant are solid components. They managed an makeshift solution, but the price was the exhaust velocity was cut in half.

    Solid Core

    Solid Core NTR
    3200° K
    Exhaust velocity (H1)16,000? m/s
    Exhaust velocity (H2)8,093 m/s
    Exhaust velocity (CH4)6,318 m/s
    Exhaust velocity (NH3)5,101 m/s
    Exhaust velocity (H2O)4,042 m/s
    Exhaust velocity (CO2)3,306 m/s
    Exhaust velocity (CO or N2)2,649 m/s

    Nuclear thermal rocket / solid core fission. It's a real simple concept. Put a nuclear reactor on top of an exhaust nozzle. Instead of running water through the reactor and into a generator, run hydrogen through it and into the nozzle. By diverting the hydrogen to a turbine generator 60 megawatts can be generated. The reactor elements have to be durable, since erosion will contaminate the exhaust with fissionable materials. The exhaust velocity limit is fixed by the melting point of the reactor.

    Solid core nuclear thermal rockets have a nominal core temperature of 2,750 K (4,490° F).

    Thrust is directly related to the thermal power of the reactor. Thermal power of 450 MWth with a specific impulse of 900 seconds will produce approximately 100,000 Newtons of thrust.

    Specific impulse (and exhaust velocity) is directly related to exhaust temperature. A temperature of 2,300 to 3,100 K will produce approximately a specific impulse of 830 to 1,000 seconds.

    As a general rule, solid core NTR have superior exhaust velocity over chemical rockets because of the low molecular weight of hydrogen propellant. Chemical LH2/LOX rockets actually run hotter than solid core NTRs, but the propellant has a much higher molecular weight.

    Hydrogen gives the best exhaust velocity, but the other propellants are given in the table since a spacecraft may be forced to re-fuel on whatever working fluids are available locally (what Jerry Pournelle calls "Wilderness re-fuelling", Robert Zubrin calls "In-situ Resource Utilization", and I call "the enlisted men get to go out and shovel whatever they can find into the propellant tanks"). For thermal drives in general, and NTR-SOLID in particular, the exhaust velocity imparted to a particular propellant by a given temperature is proportional to 1 / sqrt( molar mass of propellant chemical ).

    The value for "hydrogen" in the table is for molecular hydrogen, i.e., H2. Atomic hydrogen would be even better, but unfortunately it tends to explode at the clank of a falling dust speck (Heinlein calls atomic hydrogen "Single-H"). Another reason to avoid hydrogen is the difficulty of storing the blasted stuff, and its annoyingly low density (Ammonia is about eight times as dense!). Exhaust velocities are listed for a realistically attainable core temperature of 3200 degrees K for the propellants Hydrogen (H2), Methane (CH4), Ammonia (NH3), Water (H2O), Carbon Dioxide (CO2), Carbon Monoxide (CO), and Nitrogen (N2).

    The exhaust velocities are larger than what one would expect given the molecular weight of the propellants because in the intense heat they break down into their components. Ammonia is nice because it breaks down into gases (Hydrogen and Nitrogen). Methane is nasty because it breaks down into Hydrogen and Carbon, the latter tends to clog the reactor with soot deposits. Water is most unhelpful since it doesn't break down much at all.

    Dr. John Schilling figures that as an order of magnitude guess, about one day of full power operation would result in enough fuel burnup to require reprocessing of the fissionable fuel elements. (meaning that while there is still plenty of fissionables in the fuel rod, enough by-products have accumulated that the clogged rod produces less and less energy) A reprocessing plant could recover 55-95% of the fuel. With reprocessing, in the long term each totally consumed kilogram of plutonium or highly enriched uranium (HEU) will yield ~1E10 newton-seconds of impulse at a specific impulse of ~1000 seconds.

    Dr. Schilling also warns that there is a minimum amount of fissionable material for a viable reactor. Figure a minimum of 50 kilograms of HEU.

    One problem with solid-core NTRs is that if the propellant is corrosive, that is, if it is oxidizing or reducing, heating it up to three thousand degrees is just going to make it more reactive. Without a protective coating, the propellant will start corroding away the interior of the reactor, which will make for some real excitement when it starts dissolving the radioactive fuel rods. What's worse, a protective coating against an oxidizing chemical is worthless against a reducing chemical, which will put a crimp in your wilderness refueling. And trying to protect against both is an engineering nightmare. Oxidizing propellants include oxygen, water, and carbon dioxide, while reducing propellants include hydrogen, ammonia, and methane. Carbon Monoxide is neither, as the carbon atom has a death-grip on the oxygen atom.

    Keep in mind that the oxidizing/reducing effect is only a problem with solid-core NTRs, not the other kinds. This is because only the solid-core NTRs have solid reactor elements exposed to the propellant (for heating).

    As of the year 2019, solid core NTR development has switch focus to using Low Enriched Uranium (LEU) fuel. Mostly because the powers-that-be are hysterically afraid of Highly Enriched Uranium (aka "Weapons-Grade") falling into the Wrong Hands. HEU drastically increases the development costs and security regulations.


    (ed note: TL;DR: solid-core nuclear-thermal-rocket single-stage-to-orbit is pointless with a thrust-to-weight ratio under 20)

    A few months ago, I spent some time describing some calculations of payload fraction that I derived to assist in the design of rocket vehicles. My motivation for getting into this type of work came about from my work on the X-33 rocket when I was an intern at the Skunk Works. I wondered how so many people could think that SSTO (single-stage-to-orbit) was a good idea when the mathematics argued against it.

    Right after I joined NASA, in early 2000, I was in a group that was looking at some really advanced concepts, and somehow or another, we got looking at using (solid core) nuclear thermal rockets for an SSTO vehicle. At first blush, the whole idea seems to make sense. Nuclear thermal rockets offer almost twice the specific impulse (Isp) of chemical rockets, and if an SSTO doesn’t have enough Isp with chemical rockets, then surely nuclear rockets must be better, right?

    Wrong. Super wrong.

    Nuclear-thermal SSTO turns out to be one of the worst ideas anyone has ever come up with, for two simple reasons: hydrogen and the lousy thrust-to-weight ratio of nuclear thermal rockets. Those are the same two reasons that make NTR lousy or marginal for nearly any other space application as well, but this post will focus on the issues surrounding NTR SSTO.

    In the case of any earth-to-orbit vehicle, you’ve got to have the thrust to get off the ground in the first place. Let’s assume that we’re dealing with a vertically-launched NTR SSTO. It has to have a vehicle thrust-to-weight ratio greater than one, and probably a fair bit better than that in the first place, just to get off the ground. So we can take those expressions that I derived before, assuming hydrogen as a propellant and the engine thrust-to-weight ratios that have been quoted by NTR proponents like Stan Borowski to quickly try to figure a payload fraction for an NTR SSTO.

    We find the propellant-mass-sensitive term λ (derivation here) assuming the liquid hydrogen has a density of 71 kg/m3, ullage of 3%, a mixture ratio of zero, and a tank structural mass factor of 10 kg/m3. This gives us a value of 0.1452 for this term.

    We find the gross-mass-sensitive term φ (derivation here) by assuming that the engine has a vacuum thrust of 15000 lbf, a weight of 5000 lbm, and vacuum thrust-to-weight of 3 to 1. I’m not even going to “ding” the engine for sea-level performance, since as we’ll see, it won’t even matter. With a vacuum T/W of 3 and the same for the sea-level T/W and an initial vehicle thrust-to-weight ratio of 1.25, and we’ll just say that the thrust structure doesn’t weigh anything either, the gross-mass-sensitive term comes out to be 0.4167.

    We’ll also ignore any recovery hardware (wings, landing gear, TPS, etc) and say all that weighs nothing. We’ll assume that the engine has a vacuum Isp of 900 seconds and that it takes 9200 m/s of delta-V to get to orbit.

    Plugging those numbers in the rocket equation gives us a mass ratio MR of 2.835 (very good!)

    and a propellant mass fraction PMF of 64.73%.

    Next we use the prop-mass-sensitive and gross-mass-sensitive terms, along with the propellant mass fraction to get the payload fraction mpayload/mgross (derivation here).

    We start out with the final mass fraction (1 – prop mass fraction PMF) of 35.27%. It doesn’t get any better than that. Then we subtract the gross-mass-sensitive term φ (41.67%). Now we could stop right here, because we’re already negative (-0.064). That is to say, even before accounting for the issues with tankage, we’re already out of performance. The engines weigh too much. But we’ll keep going and subtract the product of the the propellant mass fraction PMF and prop-mass-sensitive term λ (0.6473*0.1452 = 0.0940) and we end up with a payload fraction of -0.1579.

    So it’s a no-go with these engines. Our payload fraction is grossly negative and we’ve got nothing. It’s clear from the magnitude of the numbers that the engine thrust-to-weight ratio is the main culprit, although the “fluffy” liquid hydrogen tanks don’t help much either.

    So what kind of engine performance would you have to have to get even a zero payload fraction? Well, I ran some rough calculations based on a variety of speculative vacuum T/W ratios for some putative NTR engine, at a few different values of specific impulse and plotted the results here:

    The graph tells the story. To get payload fractions of zero (a launch vehicle of infinite size) you have to have a T/W at 900 sec Isp of over 10 (red circle). That’s more than three times the T/W that Stan Borowski projects for his sporty 15K NTR design, which he says will have a T/W of 3. So if you think that Stan or others can design an NTR that only weighs a third of what he thinks it will weigh, then you can dream about an NTR SSTO of infinite mass.

    As for me, I’ve thought for some time that NTR was a really bad idea for almost every application for which it is considered. The SSTO application is probably the worst. (referring back to his earlier post, where he showed that for current NTR T/W ratios, and LH2 tank masses, NTRs only get a modest improvement in payload over a comparable LOX/LH2 system, but at much higher development cost.)

    (ed note: Keeping in mind that a closed-cycle gas core NTR can have an Isp of 3,000 seconds or so.)


    Dangerous radiation. Overstuffed pantries. Cabin fever. NASA could sidestep many of the impediments to a Mars mission if they could just get there faster. But sluggish chemical rockets aren't cutting it — and to find what comes next, one group of engineers is rebooting research into an engine last fired in 1972.

    The energy liberated by burning chemical fuel brought astronauts to the moon, but that rocket science makes for a long trip to Mars. And although search for a fission-based shortcut dates back to the 1950s, such engines have never flown. In August, NASA boosted those efforts when the agency announced an $18.8-million-dollar contract with nuclear company BWXT to design fuel and a reactor suitable for nuclear thermal propulsion (NTP), a rocket technology that could jumpstart a new era of space exploration.

    "The strengths with NTP are the ability to do the very fast round trip [to Mars], the ability to abort even if you're 2 to 3 months into the missions, the overall architectural robustness, and also the growth potential to even more advanced systems," Michael Houts, principal investigator for the NTP project at NASA's Marshall Space Flight Center, told [Superfast Space Propulsion Concepts (Images)]

    Superior gas mileage

    NTP rockets would pull all that off by offering about twice the bang for the buck that chemical rockets do (exhaust velocity of 8,000 m/s as opposed to 4,400 m/s). Rather than burning fuel with oxygen, a nuclear fission reactor would serve as a powerful furnace, heating liquid hydrogen and expelling the resulting gas for thrust. How much oomph a rocket gets from its fuel depends largely on how fast it can hurl particles out the back, which in turn hinges on their mass. And NTP's single or double hydrogen atoms would be up to a dozen times lighter than chemical rocket outputs. 

    That atomic bean counting could add up to significant time savings. "Nuclear thermal propulsion can enable you to get to Mars faster, on the order of twice as fast," said Vishal Patel, a researcher involved in subcontract work for BWXT at the Ultra Safe Nuclear Corp. in Los Alamos, New Mexico. "We're looking at nice 3- to 4-month transit times."

    New tricks with an old technology

    Unlike truly exotic propulsion proposals using antimatter or nuclear fusion, researchers have long considered nuclear fission rockets technologically feasible. Concrete development began with the Atomic Energy Commission's Project Rover in 1955 — three years before NASA's founding — and continued with the NERVA rocket prototype, which fired for nearly 2 hours straight during ground tests before budget cuts ended development in 1972. 

    By then, NASA had already canceled Apollo 18 through 20, as well as Saturn V rocket production. When Mars plans followed suit, the multibillion-dollar NERVA project lost its main purpose, Houts said. The technology saw a brief revival in the late '80s and early '90s with the Space Nuclear Thermal Propulsion (SNTP) program, which also ran out of funding before flight testing.  

    But now, with interest turning back toward Mars, past research is finding new life in current projects. 

    "The key thing is, [the NERVA rocket] was extremely well documented," said John Helmey, project manager for BWXT's NTP project. "We aren't starting from scratch. We're building upon really good work that was done back in that time frame," he told Over the course of the contract, which extends through 2019, BWXT will develop conceptual designs focusing on fuel elements and the reactor core.

    Three main challenges distinguish modern efforts from the legacy research.

    Nuclear-testing rules have changed, said Jonathon Witter, BWXT NTP project chief engineer. The potential for trace levels of radioactivity in the engine exhaust means that engineers can no longer let clouds of hydrogen gas billow into the atmosphere. Instead, BWXT plans to test a trick developed at NASA's Stennis Space Center and combust the hydrogen gas with oxygen to make easy-to-catch water. Early, small-scale demonstrations will use non-nuclear hydrogen gas to test this exhaust-capturing method, but water from future nuclear tests could be decontaminated with off-the-shelf technology.

    Engineers are also redesigning the fuel elements with new materials surrounding the uranium fuel particles, according to Witter. Rocket efficiency depends on temperature too, and BWXT expects that a ceramic and tungsten composite will allow for better operation at higher temperatures.

    What's more, NERVA ran on 90 percent highly enriched uranium that would today qualify as weapons-grade. But because the fission process throws off more than enough heat, those levels are overkill, Patel said. BWXT's designs will harness material enriched to just below 20 percent (19.75%), putting it in the less-tightly regulated low enriched uranium (LEU) category. On top of allowing safer reactors, the modest levels of fissionable material could open the door to more public-private partnerships. (I'm sure that got Elon Musk's attention)

    "The LEU thing really enables the idea that non-governmental entities can get in on this," Patel said. "It's potentially game changing."

    But extensive design and many years of testing separate NTP's on-paper potential from single-season jaunts to the Red Planet, and nuclear space technology's history of false starts makes it a long shot for NASA's early Mars missions, currently scheduled for the 2030s. "It's one of several advanced propulsion options, Houts said. "There're a lot of good options that use chemical systems, and options that use electric propulsion."

    Scott Hall, a developer of one such electric propulsion prototype that recently broke records at the University of Michigan, says he'd love to see any of these technologies get into space, but doesn't think it'll happen soon. 

    "Optimistically, it will be 15 years," Hall said of his high-powered ion thruster, "and realistically it's probably more like 50… The process is just moving so slowly, and I imagine the nuclear guys are in a similar boat." 

    But whether it takes one decade or 10, Houts thinks nuclear technology could transform space exploration. He cites Martian power plants and the possibility of spacecraft that refuel from naturally occurring resources like water or methane as examples of far-off possibilities. 

    "What we're talking about is a first generation system. The systems beyond that could have extremely advanced capabilities," he said.

    Nasty Methane Carbon Build-up

         Rip started to announce his name, rank, and the fact that he was reporting as ordered. Commander O’Brine brushed his words aside and stated flatly, “You’re a Planeteer. I don’t like Planeteers.”
         Rip didn’t know what to say, so he kept still. But sharp anger was rising inside of him.
         O’Brine went on, “Instructions say I’m to hand you your orders en-route. They don’t say when. I’ll decide that. Until I do decide, I have a job for you and your men. Do you know anything about nuclear physics?”
         Rip’s eyes narrowed. He said cautiously, “A little, sir.”
         “I’ll assume you know nothing. Foster, the designation SCN means Space Cruiser, Nuclear. This ship is powered by a nuclear reactor. In other words, an atomic pile. You’ve heard of one?”
         Rip controlled his voice, but his red hair stood on end with anger. O’Brine was being deliberately insulting. This was stuff any Planeteer recruit knew. “I’ve heard, sir.”
         “Fine. It’s more than I had expected. Well, Foster, a nuclear reactor produces heat. Great heat. We use that heat to turn a chemical called methane into its component parts. Methane is known as marsh gas, Foster. I wouldn’t expect a Planeteer to know that. It is composed of carbon and hydrogen. When We pump it into the heat coils of the reactor, it breaks down and creates a gas that burns and drives us through space. But that isn’t all it does.”
         Rip had an idea What was coming, and he didn’t like it. Nor did he like Commander O’Brine. It was not until much later that he learned that O’Brine had been on his way to Terra to see his family for the first time in four years when the cruiser’s orders were changed. To the commander, whose assignments had been made necessary by the needs of the Special Order Squadrons, it was too much. So he took his disappointment out on the nearest Planeteer, who happened to be Rip.
         “The gases go through tubes,” O’Brine went on. “A little nuclear material also leaks into the tubes. The tubes get coated With carbon, Foster. They also get coated with nuclear fuel. We use thorium. Thorium is radioactive. I won’t give you a lecture on radioactivity, Foster. But thorium mostly gives off the kind of radiation known as alpha particles. Alpha is not dangerous unless breathed or eaten. It won’t go through clothes or skin. But when mixed with fine carbon, thorium alpha contamination makes a mess. It’s a dirty mess, Foster. So dirty that I don’t want my spacemen to fool with it.

    (ed note: now in a real solid-core NTR, nuclear fuel leaking from the reactor elements is a major malfunction)

         “I want you to take care of it instead,” O’Brine said. “You and your men. The deputy commander will assign you to a squadron. Settle in, then draw equipment from the supply room and get going. When I want to talk to you again, I’ll call for you. Now blast off, Lieutenant, and rake that radiation. Rake it clean.”
         Rip forced a bright and friendly smile. “Yes, sir,” he said sweetly. “We’ll rake it so clean you can see your face in it, sir.” He paused, then added politely, “If you don’t mind looking at your face, sir—to see how clean the tubes are, I mean.”
         Rip turned and got out of there.
         Koa was waiting in the passageway outside. Rip told him what had happened, mimicking O’Brine’s Irish accent.
         The sergeant-major shook his head sadly. “This is what I meant, Lieutenant. Cruisers don’t clean their tubes more’n once in ten accelerations. The commander is just thinking up dirty work for us to do, like I said.”
         “Never mind,” Rip told him. “Let’s find our squadron and get settled, then draw some protective clothing and equipment. We’ll clean his tubes for him. Our turn will come later.”
         He remembered the last thing Joe Barris had said, only a few hours before. “Joe was right,” he thought. “To ourselves we’re supermen, but to the spacemen we’re just simps.” Evidently O’Brine was the kind of space officer who ate Planeteers for breakfast.
         Rip thought of the way the commander had turned red with rage at that crack about his face, and resolved, “He may eat me for breakfast, but I’ll try to be a good, tough mouthful!”
         Commander O’Brine had not exaggerated. The residue of carbon and thorium on the blast tube walls was stubborn, dirty, and penetrating. It was caked on in a solid sheet, but when scraped, it broke up into fine powder.
         The Planeteers wore coveralls, gloves, and face masks with respirators, but that didn’t prevent the stuff from sifting through onto their bodies. Rip, who directed the work and kept track of the radiation with a gamma-beta ion chamber and an alpha proportional counter, knew they would have to undergo personal decontamination.

    (ed note: in a real rocket, the tubes would be in vacuum, so the crew would need space suits. The tubes would also be close to the reactor. The reactor is not very radioactive if it is shut down, except for neutron activation.)

         He took a reading on the ion chamber. Only a few milliroentgens of beta and gamma radiation. That was the dangerous kind, because both beta particles and gamma rays could penetrate clothing and skin. But the Planeteers wouldn’t get enough of a dose to do any harm at all. The alpha count was high, but so long as they didn’t breathe any of the dust it was not dangerous.
         The Scorpius had six tubes. Rip divided the Planeteers into two squads, one under his direction and one under Koa’s. Each tube took a couple of hours’ hard Work. Several times during the cleaning the men would leave the tube and go into the main mixing chamber while the tube was blasted with live steam to throw the stuff they had scraped off out into space.
         Each squad was on its last tube when a spaceman arrived. He saluted Rip. “Sir, the safety officer says to secure the tubes.”
         That could mean only one thing: deceleration. Rip rounded up his men. “We’re finished. The safety officer passed the word to secure the tubes, which means we’re going to decelerate.” He smiled grimly. “You all know they gave us this job just out of pure love for the Planeteers. So remember it when you go through the control room to the decontamination chamber.”
         The Planeteers nodded enthusiastically.
         Rip led the way from the mixing chamber through the heavy safety door into the engine control room. His entrance was met with poorly concealed grins by the spacemen.
         Halfway across the room Rip turned suddenly and into Sergeant major Koa. Koa fell to the deck arms flailing for balance—but flailing against his protective clothing. The other Planeteers rushed to pick him up, and somehow all their arms and hands beat against each other.
         The protective clothing was saturated with fine dust. It rose from them in a choking cloud, was picked up, and dispersed by the ventilating system. It was contaminated dust. The automatic radiation safety equipment filled the ship with an earsplitting buzz of warning. Spacemen clapped emergency respirators to their faces and spoke unkindly of Rip’s Planeteers in the saltiest space language they could think of.
         Rip and his men picked up Koa and continued the march to the decontamination room, grinning under their respirators at the consternation around them. There was no danger to the spacemen since they had clapped on respirators the moment the warning sounded. But even a little contamination meant the whole ship had to be gone over with instruments, and the ventilating system would have to be cleaned.
         The deputy commander met Rip at the door of the radiation room. Above the respirator, his face looked furious.
         “Lieutenant,” he bellowed, “haven’t you any more sense than to bring contaminated clothing into the engine control room?”
         Rip was sorry the deputy commander couldn’t see him grinning under his respirator. He said innocently, “No, sir. I haven’t any more sense than that.”
         The deputy grated, “I’ll have you up before the Discipline Board for this.”
         Rip was enjoying himself thoroughly. “I don’t think so, sir. The regulations are very clear. They say, ‘It is the responsibility of the safety officer to insure compliance with all safety regulations both by complete instructions to personnel and personal supervision.’ Your safety officer didn’t instruct us and he didn’t supervise us. You better run him up before the Board.”
         The deputy commander made harsh sounds into his respirator. Rip had him, and he knew it. “He thought even a stupid Planeteer had sense enough to obey radiation safety rules,” he yelled.
         “He was wrong,” Rip said gently. Then, just to make himself perfectly clear, he added, “Commander O’Brine was within his rights when he made us rake radiation. But he forgot one thing. Planeteers know the regulations, too. Excuse me, sir. I have to get my men decontaminated.”
         Inside the decontamination chamber, the Planeteers took off their masks and faced Rip with admiring grins. For a moment he grinned back, feeling pretty good. He had held his own with the spacemen, and he sensed that his men liked him.
         “All right,” he said briskly. “Strip down and get into the showers.”
         In a few moments they were all standing under the chemically treated water, washing off the contaminated dust. Rip paid special attention to his hair, because that was where the dust was most likely to stick. He had it well lathered when the Water suddenly cut off. At the same moment, the cruiser shuddered slightly as control blasts stopped its spinning and left them all weightless. Rip saw instantly what had happened. He called, “All right, men. Down on the floor.”
         The Planeteers instantly slid to the shower deck. In a few seconds the pressure of deceleration pushed at them.
         “I like spacemen,” Rip said wryly. “They wait until just the right moment before they cut the water and decelerate. Now we’re stuck in our birthday suits until we land—wherever that may be.”

    From Rip Foster Rides the Gray Planet by Blake Savage {Harold Leland Goodwin} (1952)


    Nuclear Engine for Rocket Vehicle Applications. The first type of NTR-SOLID propulsion systems. It used reactor fuel rods surrounded by a neutron reflector. Unfortunately its thrust to weight ratio is less than one, so no lift-offs with this rocket. The trouble was inadequate propellant mass flow, the result of trying to squeeze too much hydrogen through too few channels in the reactor.

    Thrust Power0.198-0.065 GW
    Exhaust velocitySee Table
    Thrust49,000 n
    Engine mass10 tonne
    T/W >1.0no
    NERVA (H2)
    Exhaust Velocity8,093 m/s
    Specific Impulse825 s
    Thrust49,000 N
    Thrust Power0.2 GW
    Mass Flow6 kg/s
    Total Engine Mass10,000 kg
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power50 kg/MW
    Resuable Nuclear Shuttle [+]
    Propulsion SystemNERVA
    Exhaust Velocity8,000 m/s
    Specific Impulse815 s
    Thrust344,000 N
    Thrust Power1.4 GW
    Mass Flow43 kg/s
    Wet Mass170,000 kg
    Dry Mass30,000 kg
    Mass Ratio5.67 m/s
    ΔV13,877 m/s
    Widmer Mars Mission [+]
    Propulsion SystemNERVA
    Exhaust Velocity8,000 m/s
    Specific Impulse815 s
    Thrust580,000 N
    Thrust Power2.3 GW
    Mass Flow72 kg/s
    Wet Mass400,000 kg
    Dry Mass150,000 kg
    Mass Ratio2.67 m/s
    ΔV7,847 m/s
    HELIOS 2nd Stage [+]
    Propulsion SystemNTR Solid
    Exhaust Velocity7,800 m/s
    Specific Impulse795 s
    Thrust981,000 N
    Thrust Power3.8 GW
    Mass Flow126 kg/s
    Wet Mass100,000 kg
    Dry Mass6,800 kg
    Mass Ratio14.71 m/s
    ΔV20,968 m/s
    Atomic V-2 [+]
    Propulsion SystemNTR Solid
    Exhaust Velocity8,980 m/s
    Specific Impulse915 s
    Thrust1,050,000 N
    Thrust Power4.7 GW
    Mass Flow117 kg/s
    Total Engine Mass4,200 kg
    Wet Mass42,000 kg
    Dry Mass17,000 kg
    Mass Ratio2.47 m/s
    ΔV8,122 m/s
    Specific Power1 kg/MW


    Pewee-class Engine
    Exhaust Velocity9,200 m/s
    Specific Impulse940 s
    Thrust111,200 N
    (25 klbf)
    Thrust Power512 MWt
    Mass Flow12.5 kg/s
    Total Engine Mass3,240 kg
    Uranium 235
    Fissle Loading0.25 g U per cm3
    Max Fuel Temp2940 K
    Fuel Element
    1.32 m
    U-235 Mass36.8 kg
    Chamber Pressure1000 psi
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Engine Length
    (inc. skirt ext)
    7.01 m
    Nozzle Skirt
    2.16 m
    Nozzle Exit Dia1.87 m
    Specific Power6.3 kg/MW
    Longest Single
    44.5 min
    Total Burn
    79.2 min
    Num Burns4

    The 25 kilo-pounds-force (25 klbf) "Pewee" solid-core nuclear thermal rocket was the smallest engine size tested during U.S. Project Rover. While small, a cluster of three is adequate for a typical Mars mission. Single engines were adequate for unmanned scientific interplanetary missions or small nuclear tugs.

    A cluster of three Pewee-class engines were selected to be used with NASA's Design Reference Architecture (DRA 5.0) Mars Mission, but later designs replaced them with a cluster of three SNRE-class.

    One source suggested that each engine would require a 2,150 kg anti-radiation shadow shield to protect the crew (6.45 metric tons total for a cluster of three), assuming an 80 meter separation between the engines and the habitat module and all the liquid hydrogen propellant tanks used as additional shielding.


    The Small Nuclear Rocket Engine (SNRE) is from the report Affordable Development and Demonstration of a Small NTR Engine and Stage: How Small is Big Enough? by Stanley Borowsky et al (2015). The scientists wanted to promote the development of a right-sized solid core nuclear thermal rocket that was as small as possible, but no smaller.

    The 111,200 N (25 klbr) "Pewee-class" from the U.S. Project Rover was the smallest Rover engine. A cluster of three was specified for the NASA DRA 5.0 reference, but Borowsky et al determined that was still a bit larger than was strictly necessary.

    They looked at a 33,000 Newton (7.5 klbr) engine which was pretty much the smallest NTR possible due to limits on nuclear criticality. There is a minimum amount of fissionable fuel for a reactor, or it just cannot support a chain reaction. But it was a bit too small to do anything useful, even in a cluster of three. About all it was good for was an unmanned robotic science mission.

    A 73,000 Newton (16.5 klbr) engine on the other hand could perform quite a few proposed missions. It hit the goldilocks zone, it was just right. Some researchers took designs from NASA's Design Reference Architecture (DRA 5.0) Mars Mission and swapped out the trio of Pewee-class engines for a trio of SNREs.

    The engine uses a graphite composite core, because that allowed them to build on the expertise from the old NERVA program.

    One source suggested that each SNRE-class would require a 2,000 kg anti-radiation shadow shield to protect the crew (six metric tons for a trio of SNREs), assuming an 80 meter separation between the engines and the habitat module.

    The criticality-limited engine has a retractable section of the nozzle, the SNRE-class engine has a nozzle skirt that folds on a hinge (see diagrams below). These are strictly for launch purposes. The spacecraft is boosted in modular parts by several flights of launch vehicle, and assembled in orbit. By retracting/folding the engine nozzle the engine's overall length is reduced enough so that the engine, the liquid hydrogen fuel tank and a small mission payload can be crammed into the launch vehicle's payload faring. Once the spacecraft is assembled, the nozzles are unretracted/unfolded and permanently latched into place.

    Criticality-limited Engine
    Exhaust Velocity8,770 m/s
    Specific Impulse894 s
    Thrust33,000 N
    (7.4 klbf)
    Thrust Power145 MWt
    Mass Flow3.8 kg/s
    Total Engine Mass1,770 kg
    Uranium 235
    Max Enrichment93% U-235 wt
    Num Fuel Elements260
    Num Tie-tube
    Fissle Loading0.6 g U per cm3
    Max Fuel Temp2736 K
    U-235 Mass27.5 kg
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power12.2 kg/MW
    Engine Length6.19 m
    Fuel Length
    Pressure Vessel
    0.877 m
    Nozzle Exit
    1.32 m
    Longest Single
    22 min
    Total Burn
    29.5 min
    Num Burns2
    SNRE-class Engine
    Exhaust Velocity8,829 m/s
    Specific Impulse900 s
    Thrust73,000 N
    (16.7 klbf)
    Thrust Power367 MWt
    Mass Flow8.4 kg/s
    Total Engine Mass2,400 kg
    Uranium 235
    Max Enrichment93% U-235 wt
    Num Fuel Elements564
    Num Tie-tube
    Fissle Loading0.6 g U per cm3
    Max Fuel Temp2,726 K
    U-235 Mass59.6 kg
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power6.5 kg/MW
    Engine Length4.46 m
    Fuel Length0.89 m
    Pressure Vessel
    0.98 m
    Nozzle Exit
    2.26 m
    Longest Single
    21.4 min
    Total Burn
    55 min
    Num Burns5


    Cermet NERVA
    Exhaust Velocity9,120 m/s
    Specific Impulse930 s
    Thrust445,267 N
    Thrust Power2.0 GW
    Mass Flow49 kg/s
    Total Engine Mass9,000 kg
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power4 kg/MW

    Cermet NTR are where the fissionable fuel elements are a composite mixture of fissionable ceramics and a metal matrix.

    The problem with the original NERVA fuel elements was the blasted things were too fragile. They were rods of uranium oxide about as strong as a fine china dish. Under the vibrations of rocket flight the rods tended to snap in two. And now you've got live radioactive nuclear fuel spewing out the exhausts like a flying Chernobyl. True the rods were clad in metal to prevent them from eroding away, but the metal less like armor and more like a foil covering. They did nothing to stop the snapping. Making the cladding any thicker caused other problems.

    Cermet made the fuel rods act more like solid bars of metal. The rods were basically (solid) foamy tungsten with fissionable uranium oxide trapped inside the bubbles. The tungsten skeleton ensured that the rods would laugh at the engine vibrations.

    Cermet NERVA
    Cermet NERVA
    Exhaust Velocity9,810 m/s
    Specific Impulse1,000 s
    Thrust134,400 N
    Thrust Power0.7 GW
    Mass Flow14 kg/s
    Total Engine Mass32,546 kg
    Frozen Flow eff.73%
    Thermal eff.96%
    Total eff.70%
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power49 kg/MW

    The NERVA (Nuclear Engine for Rocket Vehicle Application) system captures the neutronic energy of a nuclear reaction using a heat exchanger cooled by water or liquid hydrogen. The exchanger uses thin foil or advanced dumbo fuel elements with cermet (ceramic-metal) substrates, jacketed by a beryllium oxide neutron reflector.

    The chamber temperature is limited to 3100K for the extended operational life of the solid fuel elements, which can be fission, fusion, or antimatter. At this temperature, the disassociation of molecular H2 to H significantly boosts specific impulse at chamber pressures below 10 atm.

    A propellant tank pressurized to 2 atm expels the LH2 coolant into the exchanger without the need for turbopumps. This open-cycle coolant is expanded through a hydrogen-cooled nozzle of refractory metal to obtain thrust.

    The efficiencies are 96% thermal, 76% frozen-flow (mainly H2 dissociation, less recombination in the nozzle), and 96% nozzle. A 940 MWth heat exchanger yields a thrust of 134 kN, and a specific impulse of 1 ksec, at a power density of 340 MW/m3.

    Altseimer, et al., “Operating Characteristics and Requirements for the NERVA Flight Engine,” AIAA Paper 70-676, June 1970.

    From HIGH FRONTIER by Philip Eklund
    Composite/Cermet Comparison


    The study author is trying to compare different types of solid-core NTR to figure out which is best. Spoiler Alert: the HEU-Cermet seems to have a slightly better performance than the others.

    All the engines were designed to have 111 kiloNewtons of thrust (25 klbf, the same as the Pewee). The first variable was Low Enriched Uranium (LEU, 2%-20% 235U) or Highly Enriched Uranium (HEU 20%-100% 235U, which includes Weapons Grade). The second variable was old-style composite fuel elements and new-style cermet elements. These gave four engines to compare:

    Case Identifiers
    HEU 93.1%com93cer93
    LEU 19.1%com20cer20

    HEU-Composite (com93) is based on the SNRE design. The SNRE had much more moderator material added in an attempt to reduce the engine mass, and to reduce the thrust.

    LEU-Composite (com20) is a reactor designed to use LEU fuel, instead of the HEU fuel used in the orginal NERVA. It is much more difficult to create a nuclear reaction with LEU fuel, but military is much happier preventing weapons-grade uranium being in civilian hands. To allow using LEU, the materials in the engine were swapped with materials that absorbed less neutrons. The amount of moderator was increased by adding more hydrogen.

    HEU-Cermet (cer93) is the standard cermet NTR. It was actually tested in the 60s and 70s. The cermet fuel elements are uranium oxide in a tungsten matrix. It has less moderating material than the NERVA.

    LEU-Cermet (cer20) is a more theoretical concept. In order to get away with using LEU, much like the com20 the engine has to be built out of materials that absorb less neutrons. You want more neutrons generating power by hitting uranium atoms and less neutrons wasted being absorbed by the engine, when LEU has fewer uranium atoms to be hit in the first place. In particular the tungsten in the cermet was to be isotopically enriched, i.e., have a higher proportion of tungsten-184 than the 31% you find in naturally occurring tungsten. Other alternatives are having the fuel elements use uranium nitride ceramic instead of the standard uranium oxide and/or using molybdenum instead of tungsten (or at least in the regions of the engine where the temperature was below molybdenum's melting point).

    Engine Descriptions

    The HEU-Composite fuel element has a width of 1.91 cm flat and 19 holes to simplify fabrication, but the 121 cm length is troubling (longest of all the four concepts). The LEU-Composite element has a wider width of 2.77 flat to mimimize neutron absorption.

    The HEU-Cermet has 91 holes for maximum specific impulse. The LEU-Cermet has only 61 holes but this only cuts the specific impulse by a few seconds. The HEU-Cermet has the shortest fuel length (64 cm) of all four concepts, which is a plus.

    The U-235 densities are interesting. The cer93 (HEU-cermet) is two orders of magnitude higher than com20 (LEU-composite): 4.560 to 0.063. Most of the design difference between the concepts can be traced back to the U-235 densities.

    All of the engines except for HEU-Cermet (cer93) use tie-tubes that contain zirconium hydride moderator. This reduced the mass but adds technical risk (i.e., it might not be possible). Zirconium hydride is prone to swelling and phase changes (melting) at various temperature under nuclear irradiation, which is a bad thing. Understand that the tie-tubes are the framework holding the engine together. After shut-down it might require the engine to keep venting huge amounts of liquid hydrogen propellant just to cool off the ZrH tie-tubes. This will drastically reduce the net specific impulse. The tie-tube hydrogen coolant is supposed to be reused, but the tie-tubes leak. Actually pretty much everything leaks hydrogen, the blasted stuff can sneak in between the atoms of the container and escape.

    ZrC40 is a 40% dense (60% porous) zirconium carbide insulator material developed during the Rover/NERVA program. It is used to prevent heat from the hot uranium fuel rods from overheating the zirconium hydride moderator in the center of the tie-rods. The fuel rod heat is supposed to be all used on the hydrogen propellant. Any leaking into the tie-rod is both lowering propellant heating efficiency and threatening the zirconium hydride with a lethal melt-down. The secret of making ZrC40 may have been lost when the NERVA program shut down, which would require redevelopment work.

    The HEU-Composite (com93) uses ZrH1.6, zirconium hydride with a H/Zr ratio of 1.62. Com20 and cer20 bump that up to a ratio of 2.0. The 2.0 ratio may be unworkable, depending upon how well it can be cooled (nominally and during transients). And depending upon how well the hydrogen coolant can be prevented from leaking out of the tie-tube, which hinges on controlling the temperature and overpressure.

    To reduce neutron capture (very important if you are trying to make do with weak LEU) the LEU-Composite (com20) uses SS-315 (marine grade stainless steel) for the tie-tube structure. Assuming the tie-tubes can be kept cool enough to prevent the steel from melting.

    The LEU-Cermet reduces neutron capture by using molybdenum that has been enriched in 96molybdenum, which has an even lower neutron capture than SS-315. It also usess W5Re instead of W25Re for the fuel coating. The lower rhenium content reduces neutron capture. The drawback is W5Re is less ductile, making it more difficult to manufacture the fuel rods and increasing the risk of rupture.

    In fact, the LEU-Cermet design has a bunch of materials enriched to remove the isotopes which are notorious for absorbing neutrons. The materials are tungsten, gadolinium, molybdenum, and rhenium. Tungsten is 98% 184W, gadolinium is 99% 160Gd, molybdenum is 98% 92Mo, and rhenium is 95% 187Re. It is unclear if these levels of enrichment could ever be practical, especially 184W. The other designs would be improved by using these enriched materials, but they can get by without them. The LEU-Cermet cannot.

    For the two cermet designs, "W10" indicates that the slat region is filled with 10% tungsten and 90% void. The designs also have an axial reflector composed of beryllium oxide on the cold end (the coolant inlet). This adds neutron reactivity, and flattens the power profile.

    Concept Performance

    Reactor Neutronics

    Each of the concepts is designed to have at least a 1% operational margin in Keff; that is, Keff > 1.01 from Beginning of Life (BOL) to End of Life (EOL). "Life" in the context of a nuclear reactor means that time from when a fresh bundle of fissionable fuel rods is installed to the time when a spent bundle of fuel rods full of nuclear poisons is removed from the reactor. When the engine is shut down after a burn (the control drums are set to quench the reactor) Keff < 0.95.

    While each concept meet the requirements, the LEU-Cermet (cer20) has so little neutron margin that it may be impossible to make a design that actually works.

    Table III above give fuel power peaking, or the factor that peak power is above average power. Which make a real difference when applying it to a reactor with very high power densities. Two kilowatts average power with a peaking power of 2.7 means the design just has to handle a peak power of 5.4 kilowatts. Easy. But if the average power is two gigawatts, designing it to handle a peak power of 5.4 GW is not trivial.

    Reactor Thermal-Hydraulics

    Table IV summarizes the core thermal-hydraulics, while tables V and VI show the temperature conditions at two potentially limiting core regions. Without a detailed thermal-structural analysis, it is hard to say which of these regions will be limiting for the entire design. The report makes a guess that it will be the peak fuel temperature location in table VI.

    The cermet engines have much higher power densities than composite cores, the cermets actually have lower fuel delta-Ts due to the smaller lattice size of the holes and higher thermal conductivity. The max fuel temperatore of 2800 K is arbitrary and probably optimistic.

    Rocket Performance

    In Table VII the power deposition (P.D.) values represent direct nuclear heating (i.e., reactor energy transferred into propellant for thrust, which we want). Loss values represent heat transfer between components (reactor energy wasted by heating up engine components, which we do not want).

    Table IX summarizes the key performance parameters of each system.

    The reactor mass includes an internal radiation shield designed to keep doses "above" (i.e., in a direction opposite to the exhaust flow) to no more than 10 MRad(Si) gamma radiation and 1×1014 n/cm2 (>100 keV). This is a radiation level that motors and turbomachinery can tolerate, but not much else. Even adding separation that much radiation will heat up the propellant tanks until they pop and give the crew lethal doses. A much heavier shield will be needed for crew safety.

    And if you cluster engines, you will need neutron isolation shields or neutrons from adjacent engines will cause nuclear flare-ups. In this case, the cermet cores will have an advantage because the side leakage is reduced by the core-internal high-Z shielding, a thicker radial reflector, and a power profile peaked closer to the radial centerline.

    The engines specific impulses (Isp) are listed (multiply by 9.81 to get exhaust velocity in m/s). Two adjustments were made.

    The "decay cooling" adjustement accounts for hydrogen flow required to prevent overheat after reactor shutdown. The engine concepts that use extra moderation (com93, com20) require significantly more cooling because of the need to keep the zirconium hydride cool.

    The "peaking change" adjustment accounts for changes in the peaking factor cause by control drum movements required for burnup reactivity effects. These include fuel depletion and fission product accumulation over the assumed 10 hours of total thrusting time between BOL and EOL. It also includes the effect of 135xenon poisoning during the (presumed) 45 minute individual burns. Xenon-135 is the most powerful known neutron-absorbing nuclear poison with a half-life of 9.2 hours. Xenon poisoning may prevent the moderated systems (com93, com20) from restarting for a burn within a day or two of prior operation. This will limit mission flexibility.


    • There is relatively little difference between the engines. Not a surprise since they all have the same fuel temperature limit
    • The net Isp of the HEU-Cermet (cer93) is about 17 seconds higher than the others, and with a higher thrust-to-weight ratio as well
    • The LEU-Cermet (cer20) has two strikes against it: peaking factors and requirement for isotopal enrichment. These could be mitigated with a higher mass design.

    NERVA Derivative

    NERVA Deriv
    Exhaust Velocity8,085 m/s
    Specific Impulse824 s
    Thrust334,061 N
    Thrust Power1.4 GW
    Mass Flow41 kg/s
    Total Engine Mass10,100 kg
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power7 kg/MW


    General Dumbo
    Thrust Power14.0-4.6 GW
    Exhaust velocitySee Table
    Thrust3,500,000 n
    Engine mass5 tonne
    T/W >1.0yes
    Dumbo (H2)
    Exhaust Velocity8,093 m/s
    Specific Impulse825 s
    Thrust3,500,000 N
    Thrust Power14.2 GW
    Mass Flow432 kg/s
    Total Engine Mass5,000 kg
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Dumbo Model A
    Engine mass0.7 tonne
    Thrust400,000 n
    Propellant mass flow52 kg/sec
    Exhaust velocity7,700 m/sec
    Engine Height0.6 m
    Engine Radius0.3 m
    Engine Volume0.2 m3
    Dumbo Model B
    Engine mass2.8 tonne
    Thrust3,560,000 n
    Propellant mass flow460 kg/sec
    Exhaust velocity7,700 m/sec
    Engine Height0.6 m
    Engine Radius1.0 m
    Engine Volume1.8 m3
    Dumbo Model C
    Engine mass2.1 tonne
    Thrust400,000 n
    Propellant mass flow48 kg/sec
    Exhaust velocity8,300 m/sec
    Engine Height0.6 m
    Engine Radius0.4 m
    Engine Volume0.3 m3

    This was a competing design to NERVA. It was shelved political decision that, (in order to cut costs on the atomic rocket projects) required both projects to use an already designed NEVRA engine nozzle. Unfortunately, said nozzle was not compatible with the DUMBO active cooling needs. Dumbo does, however, have a far superior mass flow to the NERVA, and thus a far superior thrust. Dumbo actually had a thrust to weight ratio greater than one. NASA still shelved DUMBO because [a] NERVA used off the shelf components and [b] they knew there was no way in heck that they could get permission for nuclear lift-off rocket so who cares if T/W < 0? You can read more about Dumbo in this document.

    Note that the "engine mass" entry for the various models does not include extras like the mass of the exhaust nozzle, mass of control drums, or mass of radiation shadow shield.

    Pebble Bed

    Pebble Bed
    Exhaust Velocity9,530 m/s
    Specific Impulse971 s
    Thrust333,617 N
    Thrust Power1.6 GW
    Mass Flow35 kg/s
    Total Engine Mass1,700 kg
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power1 kg/MW

    Particle bed / nuclear thermal rocket AKA fluidized-bed, dust-bed, or rotating-bed reactor. In the particle-bed reactor, the nuclear fuel is in the form of a particulate bed through which the working fluid is pumped. This permits operation at a higher temperature than the solid-core reactor by reducing the fuel strength requirements . The core of the reactor is rotated (approximately 3000 rpm) about its longitudinal axis such that the fuel bed is centrifuged against the inner surface of a cylindrical wall through which hydrogen gas is injected. This rotating bed reactor has the advantage that the radioactive particle core can be dumped at the end of an operational cycle and recharged prior to a subsequent burn, thus eliminating the need for decay heat removal, minimizing shielding requirements, and simplifying maintenance and refurbishment operations.

    Project Timberwind

    Project Timberwind was started in President Reagan infamous Strategic Defense Initiative ("Star Wars"). It was later transferred to the Air Force Space Nuclear Thermal Propulsion (SNTP) program. The project was cancelled by President William Clinton.

    NTR Comparison
    Engine Mass6,803 kg1,500 kg
    Thrust (Vac)333.6 kN392.8 kN
    Specific Impulse850 s1,000 s
    Burn Time1,200 s449 s
    T/W530 !!!

    The idea was to make a nuclear-powered interceptor to destroy incoming Soviet ICBMs. The Timberwind NTR upper stage would have to make the NERVA engine look like a child's toy, with huge specific impulse and an outrageously high thrust-to-weight ratio. The project managers babbled about advances in high-temperature metals, computer modelling and nuclear engineering in general justifying suspiciously too-good-to-be-true performance. It was based on the pebble-bed concept.

    Diameter4.25 m2.03 m8.70 m
    Thrust (Vac)392.8 kN735.5 kN2,451.6 kN
    Specific Impulse1,000 s1,000 s1,000 s
    Engine Mass1,500 kg2,500 kg8,300 kg
    Burn Time449 s357 s493 s

    Pulsed Solid-core NTR

    The pulsed nuclear thermal rocket is a type of solid-core nuclear thermal rocket concept developed at the Polytechnic University of Catalonia, Spain and presented at the 2016 AIAA/SAE/ASEE Propulsion Conference. It isn't a torchship but it is heading in that direction. Thanks to Isaac Kuo for bringing this to my attention.

    As previously mentioned, solid core nuclear thermal rockets have to stay under the temperature at which the nuclear reactor core melts. Having your engine go all China Syndrome on you and shooting out what's left of the exhaust nozzle in a deadly radioactive spray of molten reactor core elements is generally considered to be a Bad Thing. But Dr Francisco Arias found a clever way to get around this by pulsing the engine like a TRIGA reactor. The engine can be used bimodally, that is, mode 1 is as a standard solid-core NTR (Dr. Arias calls this "stationary mode"), and mode 2 is pulsed mode.

    Pulse mode can be used two ways:

    Direct Thrust Amplification: Garden variety solid core NTRs can increase their thrust by shifting gears. You turn up the propellant mass flow. But since the reactor's energy has to be divided up to service more propellant per second, each kilogram of propellant gets less energy, so the exhaust velocity and specific impulse goes down.

    But if you shift to pulse mode along with increased propellant mass flow, the reactor's effective energy output increases. So you can arrange matters in such a way that each kilogram of propellant still gets the same share of energy. Bottom line: the thrust increases but the specific impulse is not degraded.

    Specific Impulse Amplification: This is really clever. For this trick you keep the propellant mass flow the same as it was.

    In a fission nuclear reactor 95% of the reactor energy comes from fission-fragments, and only 5% come from prompt neutrons. In a conventional solid-core NTR the propellant is not exposed to enough neutrons to get any measurable energy from them. All the energy comes from fission fragments.

    But in pulse mode, that 5% energy from neutrons could be higher than the 95% fission-fragment energy in stationary mode. The difference is that fission fragment energy heats the reactor and reactor heat gives energy to the propellant. And if the reactor heats too much it melts. But neutron energy does not heat the reactor, it passes through and directly heats the propellant.

    The end result is that in pulse mode, you can actually make the propellant hotter than the reactor. Which means a much higher specific impulse than a conventional solid-core NTR which running hot enough to be right on the edge of melting.

    Thermodynamics will not allow heat energy to pass from something colder to something hotter, so it cannot make the propellant hotter than the reactor. But in this case we are heating the propellant with neutron kinetic energy, which has zippity-do-dah to do with thermodynamics.

    The drawback of course is that the 95% fission-fragment energy is increased as well as the neutron energy. The important point is by using pulsing you can use an auxiliary cooling system to cool the reactor off before the blasted thing melts, unlike a conventional NTR.

    Apparently Dr. Arias' paper claims the pulsed NTR can have a higher specific impulse than a fission fragment engine. I am no rocket scientist but I find that difficult to believe. Fission fragment can have a specific impulse on the order of 1,000,000 seconds.

    How Does It Work?

    TRIGA reactor have what is called a large, prompt negative fuel temperature coefficient of reactivity. Translation: as the nuclear fuel elements heat up they stop working. It automatically turns itself off if it gets too hot. Technical term is "quenching."

    Which means you can overload it in pulses. The TRIGA is designed for a steady power level of 100 watts but you can pulse the blasted thing up to 22,000 freaking megawatts. It automatically shuts off after one-twentieth of a second, quickly enough so the coolant system can handle the waste heat pulse.

    Amplification Factor

    The amount of amplification of thrust or specific impulse requires the value of N, or energy ratio between the pulsed mode and the stationary mode (pulsed mode energy divided by stationary mode energy). This can be calculated by the formidable equation

    ΔT is the temperature increase during a pulse (in Kelvin), t is the residence time of the propellant in the reactor (seconds), and [ ΔT/t ] is the quench rate (K/sec). ΔT will probably be about 103 K (assuming propellant velocity of hundreds of meters per second and chambers about one meter long), t will probably be from 10-3 sec to 10-2 sec. This means [ ΔT/t ] will be about 105 to 106 K/s.

    I'm not going to explain the other variables, you can read about them here.

    Be that as it may, Wikipedia states that if you use standard reactor fuels like MOX fuel or Uranium dioxide, fuel heat capacity ≅ 300J/(mol ⋅ K), fuel thermal conductivity ≅ 6W/(K ⋅ m2), fuel density of ≅ 104kg/(m3), cylindrical fuel radius of ≅ 10-2m and a fuel temperature drop from centerline to cladding edge of 600K then:

    N ≅ 6×10-3 * [ ΔT/t ]

    This boils down to N being between 600 and 6,000.

    Direct Thrust Amplification Details

    Thrust power is:

    Fp = (F * Ve ) / 2

    Thrust is:

    F = mDot * Ve

    Specific Impulse is:

    Isp = Ve / g0


    Fp = Thrust Power (w)
    F = Thrust (N)
    Ve = Exhaust Velocity (m/s)
    mDot = Propellant Mass Flow (kg/s)
    Isp = Specific Impulse (s)
    g0 = acceleration due to gravity (9.81 m/s2)

    With a conventional solid NTR, thrust power is a constant. So if you wanted to increase the thrust by, for instance 5 time, you have to increase the propellant mass flow by 52 = 25 times and decrease the exhaust velocity by 1/5 = 0.2 times. Which decreases the specific impulse 0.2 times.

    But a pulsed NTR can increase thrust power. So if you want to increase the thrust by 5 times, you increase the thrust power by 5 times, the propellant mass flow five times, and keep the exhaust velocity and specific impulse the same.

    The limit on the increase in thrust power is N.

    Specific Impulse Amplification Details

    If in pulse mode the amplification factor is N, then the amplified specific impulse is:

    IspPulse = IspS * sqrt[ (fn * N) + 1]


    IspPulse = Specific Impulse in Pulse Mode
    IspS = Specific Impulse in Stationary Mode
    fn = fraction of the prompt neutrons (0.05)
    N = energy amplification by pulsing the reactor
    sqrt[x] = square root of x

    So if N is between 600 and 6,000, the specific impulse will increase by a factor of 5.57 to 17.35. With a basic NERVA having a specific impulse of about 800 seconds, a pulsed version would have instead 4,460 to 13,880 seconds!

    Russian Twisted Ribbon

    These are from Russian Nuclear Rocket Engine Design for Mars Exploration by Vadim Zakirov and Vladimir Pavshook. The unique "twisted ribbon" fuel elements were developed in the Soviet Union, and continued development in Russia. The twisted ribbon surface-to-volume ratio is 2.6 times higher than that of the US NERVA fuel elements, which enhances the heat transfer between fuel and propellant.

    The prototype RD-0140 engine was a pure rocket engine, while the nuclear power and propulsion system (NPPS) is a Bi-Modal NTR acting as an electrical power generator in between thrust periods. A spacecraft designed for a Mars mission would have three or four NPPS engines.

    Twisted Ribbon Engines
    Thrust (vac) (kN)35.2868
    PropellantH2 + HexaneH2
    Propellant Mass Flow (kg/s)~4~7.1
    Specific Impulse (vac) (s)~900~920
    Core outlet temparture (K)3,0002,800 to 2,900
    Chamber Pressure (105 Pa)7060
    U235 enrichment (%)9090
    Fuel Composition(U,Nb,Zn)CU-Zr-C-N
    Fuel Element FormTwisted ribbonTwisted ribbon
    Generated electrical power (kW)N/A50
    Working fluid for power loop
    (% by mass)
    N/A93% Xe + 7% He
    Max temp for power loop (K)N/A1,500
    Max press for power loop (105 Pa)N/A9
    Working fluid flow rate (kg/s)N/A1.2
    Thermal power - propulsion mode (MW)196340
    Thermal power - power mode (MW)N/A0.098
    Core length (mm)800700
    Core diameter (mm)500515
    Engine length (mm)3,700No Data
    Engine diameter (mm)1,200No Data
    Lifetime - propulsion mode (h)15
    Lifetime - power mode (yr)N/A2
    Mass (kg)2,000*1,800**

    N/A = not applicable. * = including radiation shield and adapter. ** = reactor mass.

    In the RD-0140 they added hexane to the liquid hydrogen propellant. Unfortunately pure hot hydrogen tended to erode the fuel elements and make the exhaust radioactive.

    Twisted Ribbon Engine
    Thrust power1,650 MW
    Exhaust velocity9,420 m/s
    Specific impulse960 s
    Thrust330,000 N
    Engine mass5,260 kg

    The CIS engine developed jointly by the US/CIS industry team of Aerojet, Energopool and B&W utilizes a heterogeneous reactor core design with hydrogen-cooled ZrH moderator and ternary carbide fuel materials. The ZrH moderator, in the form of close-packed rods, is located between reactor fuel assemblies and is very efficient in minimizing the inventory of fissile material in the reactor core.

    The CIS fuel assembly (shown in Figure 6) is an axial flow design and contains a series of stacked 45 mm diameter bundles of thin (~1 mm) "twisted ribbon" fuel elements approximately 2 mm in width by 100 mm in length.

    The "fueled length" and power output from each assembly is determined by specifying the engine thrust level and hydrogen exhaust temperature (or desired Isp).

    For the 75 klbf (330,000 N) CIS engine design point indicated in Figure 4, 102 fuel assemblies (each containing 10 fuel bundles) produce ~1650 MWt with a Isp of ~960 s.

    For a 15 klbf (67,000 N) engine, 34 fuel assemblies (with 6 fuel bundles each) are used to generate the required 340 MWt of reactor power at the same Isp.

    The fuel material in each "twisted ribbon" element is composed of a solid solution of uranium, zirconium and niobium ceramic carbides having a maximum operating temperature expected to be about 3200 K. The fuel composition along the fuel assembly length is tailored to provide increased power generation where the propellant temperature is low and reduced power output near the bottom of the fuel assembly where the propellant is nearing its exhaust temperature design limit. In the present CIS design a value of 2900 K has been selected to provide a robust temperature margin. During reactor tests, hydrogen exhaust temperatures of 3100 K for over one hour and 2000 K for 2000 hours were demonstrated in the CIS.

    At 2900 K, an engine lifetime of ~4.5 hours is predicted.

    The Aerojet, Energopool, B&W NTR design utilizes a dual turbopump, recuperated expander cycle. Hydrogen flowing from each pump is split with ~84% of the flow going to a combination recuperator/gamma radiation shield and the remaining 16% used to cool the nozzle. The recuperator/shield, located at the top of the engine, provides all of the necessary turbine drive power. The turbine exhaust cools the reactor pressure vessel and is then merged with the nozzle coolant to cool the moderator and reflector regions of the engine. The coolant then passes through borated ZrH and lithium hydride (LiH) neutron shields located within the pressure vessel between the reactor core and the recuperator/gamma shield, before returning to the recuperator where it heats the pump discharge flow. Exiting the recuperator the cooled hydrogen is then routed to the core fuel assemblies where it is heated to 2900 K.

    The 75 klbf (330,000 N) CIS engine design point has a chamber pressure of 2000 psia (14,000 kpa), a nozzle area ratio of 300 to 1, and a 110% bell length nozzle resulting in a Isp of ~960 s.

    (ed note: from the chart, the 75 klbf CIS engine has a thrust-to-weight ratio of 6.4. If my slide rule is not lying to me, that means the engine has a mass of 5,260 kilograms)

    The same pressure and nozzle conditions were maintained for the 15 (67,000), 25 (110,000) and 50 klbf (220,000 N) engine design points with the resulting weight scaling indicated in Figure 4.

    The approximate engine lengths for the 15 (67,000), 25 (110,000), 50 (220,000) and 75 klbf (330,000 N) CIS engines are 4.3 m, 5.2 m, 6.5 m, and 7.6 m, respectively.

    Low Pressure NTR

    Low Pressure NTR
    Engine Mass835 kg
    Full Thrust49,000 newtons
    Full T/W6.0
    Full Isp1,210 sec
    Single-H Thrust9,800 newtons
    Single-H T/W1.2
    Single-H Isp1,350 sec

    This is from Low Pressure Nuclear Thermal Rocket (LPNTR) concept (1991)

    This is a theoretical concept, but it has enough impressive advantages over conventional solid-core NTRs that it is well worth looking into. The engine has a specific impulse of up to 1,350 seconds (exhaust velocity 13,200 m/s) which is virtually the theoretical maximum for solid-core NTR. It also is very lightweight plus much more reliable. The latter is due to the absence of certain heavy and fault-prone components (those with moving parts) required for solid-core.

    Solid-core NTRs commonly use liquid hydrogen as propellant, since that is the propellant with the sweet spot of low molecular weight and convenience. The lower the molecular weight, the higher the specific impulse and exhaust velocity.

    There is one propellant with an even lower molecular weight, but it is anything but convenient. Monatomic hydrogen has half the molecular weight of molecular hydrogen so it has a much higher performance. A pity it explodes like a bomb if you give it a stern look. In his novels Robert Heinlein calls monatomic hydrogen "Single-H", and handwaves really hard that future engineers will figure out some way to stablize the dire stuff. Sorry Mr. Heinlein, we need a real-world solution here.

    Heating molecular hydrogen to above 3,000 Kelvin will dissociate it into single-H. Sadly at the high pressures commonly used in solid-core reactors, the temperature and the propellant mass flow would combine into a heat flux high enough to destroy the reactor. Remember the difference between heat and temperature: temperature is an interesting number but it is the heat joules that ruin the reactor.

    Dr. Ramsthaler said "Ah, but what if we designed the engine to use low pressure?" Then we can make single-H at a heat flux low enough for the reactor to survive, allowing our specific impulse will climb to amazing levels. A standard NERVA has an engine pressure of 31 bar (450 pounds force per square inch), the LPNTR only has a pressure of 1 bar (14.5 psia). This means the LPNTR has a heat flux that is 50-to-one less than the NERVA.

    The drawback is the low pressure will drastically reduce the propellant mass flow, which reduces the thrust (because thrust = propellant mass flow times exhaust velocity). This problem can be addressed with clever engineering. Dr. Ramsthaler thinks it is possible to push the engine up to a thrust-to-weight ratio of 1.2. The Monatomic-H MITEE tries the same low-pressure trick, but only at a thrust-to-weight ratio of 1.0.

    Everything comes at a cost. The engine can do a T/W ratio of 6.0 at full thrust, but this means the specific impulse is only 1,210 seconds. If you shift it into temperatures that allow dissociation to create Single-H, the T/W ratio is only 1.2 but the Single-H makes a specific impulse of 1,350 seconds. So the engine has two gears.

    In addition, a low pressure engine means it does not need turbopumps to create high pressure. Turbopumps are penalty-weight, turbopumbs need complicated plumbing to supply the energy needed to spin the little darling, and turbopumps contain several points of mechanical failure with all their moving parts. Good riddance to bad rubbish. The natural propellant tank pressure is enough for the LPNTR to operate.

    Also the low heat flux means the engine only needs an exhaust nozzle that is very short compared to a NERVA. 50-to-one less than the NERVA, remember?

    Dr. Ramsthaler's secret is a reactor with a radial outflow core: it maximizes propellant mass flow at low pressure but high temperature. Remember:

    • High temperature is needed to make Single-H and crank up the specific impulse to 11, er, ah, 1,350 seconds
    • Low pressure counteracts the high temperature so the heat level is not high enough to melt the reactor
    • Maximizing propellant mass flow counteracts the low pressure so the thrust-to-weight ratio is at least 6.0

    For standard NERVA and related solid-core NTRs, at low pressure the critical flow is where the propellant exits the core. The propellant enters the top of the cylindrical core, is heated inside the core, and exits the core at the bottom. Then it enters the exhaust nozzle.

    Dr. Ramsthaler's design uses a spherical core. The propellant enters the center of the core, is heated inside the core, and exits the core from its surface. Given the 120 flow outlet holes on the surface, the engine has almost 50% flow area at the exit of the core.

    The design can accommodate almost any kind of nuclear fuel elements: pebbles, plates, whatever.

    Safety and reliabily was Dr. Ramsthaler's primary goal. But his solution to control of the nuclear reactor raises eyebrows.

    Conventional NERVA engines use control drums to control the criticality in the nuclear reactor. Spin the drums so the neutron reflector face the nuclear fuel elements and the reactor fires up. Spin the drums so the neutron poison faces the fuel elements and the reactor shuts down like a blown-out match.

    As it turns out, the liquid hydrogen propellant is a pretty good neutron moderator all by itself. The spacecraft engineer has to be careful about feeding propellant into a dry hot reactor. Otherwise neutron transients will build into full-fledged runaway nuclear oscillations and your reactor will go all Chernobyl on you. The addition of the moderator changes the nuclear characteristics of the reactor.

    Anyway Dr. Ramsthaler looked at the way the propellant altered the reactor behaviour and wondered if careful propellant control could replace the control drums. Control drums are penalty-weight, control drums require electricity, and control drums contain several points of mechanical failure with all their moving parts. Using propellant to control the reactor would happily reduce the engine mass even more, and increase the engine reliabilty.

    The hydrogen propellant is injected into the center of the spherical core, remember? This turns out to be the perfect location for the hydrogen to moderate the neutrons flux, where the neutrons are thickest. The hydrogen turns worthless fast neutrons into reactor-grade thermal neutrons which maintain the fission chain reaction.

    The dry reactor just sits there, its nuclear characteristics are such that no chain reaction can happen. But as soon as the liquid hydrogen fills the center, the reactor goes critical and starts generating large amounts of thermal energy by the miracle of nuclear fission.

    But just in case the reaction gets out of hand, there is a rod of neutron poison that can be slammed into the center of the core to scram the engine.

    Dr. Ramsthaler figures with such low engine mass, the spacecraft could afford to have seven engines. This would allow thrust vectoring by throttling engines instead of the mechanical nightmare of gimbaled engines. All together now: engine gimbals are are penalty-weight, engine gimbals require hydraulics, and engine gimbals contain several points of mechanical failure with all their moving parts. Get rid of them.

    Rob Davidoff points out that the above gimbal-less scheme will do yaw and pitch thrust vectoring just fine. But it is incapable of performing roll vectoring. A spacecraft using such a scheme will have to rely upon its reaction control system (attitude jets) for rolls.

    In addition, a cluster of seven engines would allow the spacecraft to lose up to two engines and still limp through the mission ("two-engine-out" capability). Instead of total mission failure and all the crew dying.

    LPNTR advantage IMEO for Mars mission
    + shield
    500 KM
    Earth Orbit
    Ref NERVA85042.68841400
    Adv NERVA92563.37131037
    3200 K
    3600 K
    3600 K
    Dual mode

    The Ref mission is a Mars mission that ends with the spacecraft in a huge ecliptic orbit around Terra. This will require lots of energy when you want to reuse the spacecraft. The 500 KM Earth Orbit mission is the Mars mission, using extra propellant and delta V to end with the spacecraft in a nice circular orbit for easy spacecraft reuse.

    You can see how the Initial Mass in Earth Orbit (IMEO) nicely drops as the engine Isp increases. And how using a dual-mode engine with the Single-H mode drops the IMEO by 77 metric tons of propellant compared to the single-mode engine.


    LANTR NERVA mode
    Exhaust Velocity9,221 m/s
    Specific Impulse940 s
    Thrust67,000 N
    Thrust Power0.3 GW
    Mass Flow7 kg/s
    RemassLiquid Hydrogen
    LANTR LOX mode
    Exhaust Velocity6,347 m/s
    Specific Impulse647 s
    Thrust184,000 N
    Thrust Power0.6 GW
    Mass Flow29 kg/s
    RemassHydrogen + Oxygen
    LANTR Both
    Uranium 235
    ReactorSolid Core
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    SpecialLow-High Gear
    Nuclear DC-X NERVA
    High Gear
    Exhaust Velocity9,810 m/s
    Specific Impulse1,000 s
    Thrust/Engine1,112,000 N
    Thrust5,560,000 N
    Thrust Power27.3 GW
    Mass Flow567 kg/s
    Specific Power7 kg/MW
    Low Gear
    Exhaust Velocity5,900 m/s
    Specific Impulse601 s
    Thrust/Engine3,336,000 N
    Thrust16,680,000 N
    Thrust Power49.2 GW
    Mass Flow2,827 kg/s
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power4 kg/MW
    Number Thrustersx5
    Total Engine Mass199,600 kg
    Uranium 235
    SpecialLow-High Gear
    Wet Mass460,000 kg

    LOX-augmented Nuclear Thermal Rocket. One of the systems that can increase thrust by lowering Isp, in other words Shifting Gears. This concept involves the use of a "conventional" hydrogen (H2) NTR with oxygen (O2) injected into the nozzle. The injected O2 acts like an "afterburner" and operates in a "reverse scramjet" mode. This makes it possible to augment (and vary) the thrust (from what would otherwise be a relatively small NTR engine) at the expense of reduced Isp

    Bi-Modal NTR

    Bimodal NTR Solid (NASA)
    Propulsion SystemNTR Solid Bimodal
    Exhaust Velocity8,980 m/s
    Specific Impulse915 s
    Thrust/Engine66,667 N
    Number Thrustersx3
    Thrust200,000 N
    Thrust Power0.9 GW
    Mass Flow22 kg/s
    Total Engine Mass6,672 kg
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Wet Mass80,000 kg
    Dry Mass26,830 kg
    Mass Ratio2.98 m/s
    ΔV9,811 m/s
    Specific Power7 kg/MW

    A useful refinement is the Bimodal NTR.

    Say your spacecraft has an honest-to-Johnny NERVA nuclear-thermal propulsion system. Typically it operates for a few minutes at a time, then sits idle for the rest of the entire mission. Before each use, one has to warm up the reactor, and after use the reactor has to be cooled down. Each of these thermal cycles puts stress on the engine. And the cooling process consists of wasting propellant, flushing it through the reactor just to cool it off instead of producing thrust.

    Meanwhile, during the rest of the mission, your spacecraft needs electricity to run life support, radio, radar, computers, and other incidental things.

    So make that reactor do double duty (that's where the "Bimodal" comes in) and kill two birds with one stone. Refer to below diagram. Basically you take a NERVA and attach a power generation unit to the side. The NERVA section is the "cryogenic H2 propellant tank", the turbopump, and the thermal propulsion unit. The power generation section is the generator, the radiator, the heat exchanger, and the compressor.

    Warm up your reactor once, do a thrust burn, stop the propellant flow and use the heat exchanger and radiator to partially cool the reactor to power generation levels, and keep the reactor warm for the rest of the mission while generating electricity for the ship.

    This allows you to get away with only one full warm/cool thermal cycle in the entire mission instead of one per burn. No propellant is wasted as coolant since the radiator cools down the reactor. The reactor supplies needed electricity. And as an added bonus, the reactor is in a constant pre-heated state. This means that in case of emergency one can power up and do a burn in a fraction of the time required by a cold reactor.

    Pretty ingenious, eh?

    Most nuclear thermal rockets do not need heat radiators because they get by with open-cycle cooling. But bimodal engines do need radiators, which makes sense with a few moment's thought. While running in power-generation mode the rocket is not thrusting. No thrust means no rocket exhaust. And no rocket exhaust means no handly plume of gas to use for open-cycle cooling. So you need a physical radiator to take care of the waste heat created by electrical power generation.

    An even further refinement is the Hybrid BNTR/EP option. This is where the electrical power output has a connection to an Ion Drive. This is a crude form of Shifting Gears: trading thrust for specific impulse/exhaust velocity. So it can do low-gear NTR thrust mode, high-gear ion-drive thrust mode, and no-thrust electricity generation mode while coasting.

    And the Pratt & Whitney company went one step better. They took the Bimodal NTR concept and merged it with the LANTR concept to make a Trimodal NTR. Called the Triton, it uses a LANTR engine to allow Shifting Gears. So it can do low-gear NTR-Afterburner thrust mode, high-gear NTR thrust mode, and no-thrust electricity generation mode while coasting.


    5.3.1 Bimodal NTR Mission Concept

    An option to Reference Mission Version 3.0 (DRM 3.0) that utilizes bimodal NTR transfer vehicles in place of the expendable NTR stages is being evaluated. A common "core" stage, used on cargo and piloted vehicles alike, is outfitted with three 15 klbf bimodal NTR engines capable of providing up to 50 kilowatts of electrical power (kWe) using any two engines The bimodal core stage is not jettisoned after the Trans-Mars Insertion (TMI) maneuver but remains with the cargo and piloted payload elements providing midcourse correction (MCC) propulsion and all necessary power during transit. Near Mars, the bimodal stage separates from the aerobraked payloads and performs its final disposal maneuvers. A key difference between Reference Mission 3.0 and the bimodal option is the absence of the aerobraked LOX/methane (CH4) Trans-Earth Insertion (TEI) stage which is replaced by an "all propulsive" bimodal NTR-powered Earth Return Vehicle (ERV) illustrated in Figure A5-7.

    The bimodal stage LH2 tank is slightly shorter than the expendable TMI stage tank at 19 meters and has a maximum LH2 propellant capacity of ~51 tons with a 3% ullage factor. A turbo-Brayton refrigeration system is located in the forward cylindrical adaptor section to eliminate LH2 boiloff during the lengthy (~4.3 year) ERV mission. A 12 kWe Brayton refrigeration system is included to remove the ~100 watts of heat flux penetrating the 2 inch MLI system in low-Earth-orbit where the highest heat flux occurs. Enclosed within the conical aft radiator section of the bimodal core stage is a closed Brayton cycle (CBC) power conversion system employing three 25 kWe Brayton rotating units (one for each bimodal reactor) which operate at ~2/3 of rated capacity, thus providing an "engine out" capability. The turbine inlet temperature of the He-Xe working gas is ~1300 K and the total system specific mass is estimated to be ~30 kg/kWe.

    A mass comparison of the bimodal NTR transfer vehicles and the Reference Mission Version 3.0 vehicles is shown in Table A5-1.

    The mass values assume a "2-perigee burn" Earth departure scenario. Overall, the bimodal approach has a lower "three-mission" initial mass than Reference Mission 3.0. In addition, the bimodal approach can reduce the operational complexity of the mission (eliminates solar array deployment/retraction) as well as eliminating the need for an aerobrake and injection stage for the Earth Return Vehicle.

    5.3.2 All Propulsive" Bimodal NTR Option Using TransHab

    Another option to the Reference Mission 3.0 under consideration is the use of a bimodal NTR stage to propulsively capture all payload elements into Mars orbit. This "all propulsive" NTR option provides the most efficient use of the bimodal engines which can supply abundant power to the spacecraft and payloads in Mars orbit for long periods. Propulsive capture into the reference "250 km by 1 sol" elliptical Mars parking orbit also makes possible the use of a standardized, reduced mass "aerodescent" shell because of the lower payload entry velocity (~4.5 km/s) encountered. From this orbit, the triconic aerobrake mass varies by only ~400 kg for a 20 ton increase in payload mass (see Section 3.3.3).

    The attractiveness of the "all propulsive" bimodal NTR option is further increased by the utilization of the lightweight, inflatable "TransHab" module discussed in Section 3.1. The substitution of TransHab for the heavier, hard-shell habitat module introduces the potential for propulsive recovery of the Earth Return Vehicle in Earth orbit and its reuse on subsequent missions. TransHab use also allows the crew to travel to and from Mars on the same bimodal transfer. In Mars orbit, the crew transfer vehicle rendezvous with the "unpiloted" habitat lander which is now delivered as a cargo element by the bimodal stage. The absence of crew from the bimodal habitat lander eliminates the need for outbound consumables and engine crew radiation shields and allows it to carry off-loaded surface habitation and science equipment previously carried on the cargo lander.

    A three-dimensional image of the bimodal transfer vehicle used on the piloted mission is shown in Figure A5-8. The TransHab is ~9.7 meters long and inflates to a diameter of ~9.5 meters. Its total mass is ~24.3 metric tons which includes the crew and their consumables. The total length and initial mass of the piloted transfer vehicle is ~54 meters and ~141 metric tons, respectively. A smaller, "in-line" propellant tank is used on the bimodal transfer vehicles that deliver the ~46 metric ton habitat and ~54 ton cargo landers into Mars orbit. The habitat and cargo transfer vehicles are ~56 meters long and have a LEO mass of ~129 metric tons and 144 metric tons, respectively.

    Dual-mode Fission
    Exhaust Velocity9,810 m/s
    Specific Impulse1,000 s
    Thrust124,700 N
    Thrust Power0.6 GW
    Mass Flow13 kg/s
    Total Engine Mass33,000 kg
    Thermal eff.94%
    Total eff.94%
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power54 kg/MW
    Thermal Electrical eff.19%
    Electrical Power60 MWe

    When struck by a thermal neutron, a fissile nuclide splits into two fragments plus energy. For example, the fission of the 235U atom produces 165 MeV of energy plus 12 MeV of neutral radiation (gammas and a couple of fast neutrons). The fast neutrons must be thermalized by a low Z moderator (a surrounding blanket of about 80 cm of D2O, Be, liquid or gas D2, or CD4), which returns enough thermal neutrons to the core to sustain the chain reaction. (Thermal neutrons diffuse through the reactor like a low pressure gas.) Alternatively, a molybdenum neutron reflector can be used. Much of a reactor’s mass is constant, regardless of power level. Therefore, nuclear power sources are more attractive at higher power levels.

    The 650 MWth system illustrated is dual mode, which can either generate electricity, or directly exhaust coolant for thrust. It uses a fast reactor with fuel tubes interspersed with cooling tubes. The coolant is lithium, which for electrical power is passed to a potassium boiler at 1650 K. The potassium vapor is passed to a static (AMTEC) or dynamic (turbine) heat engine for power generation (60 MWe), or heats hydrogen in a heat exchanger for thrust (125 kN at a specific impulse of 1 ks). The thermal efficiency is 19% if closed-cycle (for power generation) or 94% if open-cycle (for thrust).

    From High Frontier by Philip Eklund
    Pebble-bed Fission Reactor
    Exhaust Velocity9,810 m/s
    Specific Impulse1,000 s
    Thrust172,700 N
    Thrust Power0.8 GW
    Mass Flow18 kg/s
    Total Engine Mass58,000 kg
    Thermal eff.94%
    Total eff.94%
    Uranium 233
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power68 kg/MW
    Electrical Power60 MWe

    This is a graphite-moderated, gas-cooled, nuclear reactor that uses spherical fuel elements called "pebbles". These tennis ball-sized pebbles are made of pyrolytic graphite (which acts as the moderator), interspersed with thousands of micro fuel particles of a fissile material (such as 235U).

    In the reactor illustrated, 360,000 pebbles are placed together to create a 120 MWth reactor. The spaces between the pebbles form the "piping" in the core for the coolant, either propellant or inert He/Xe gas.

    The design illustrated can is dual mode. It can operate either as a generator for 60 MWe of electricity, or act as a solid-core thruster using hydrogen propellant/coolant expelled at a specific impulse of 1 ksec. When used as a thruster, it offers a slight increase in specific impulse but significant acceleration benefits over traditional fission reactors. Moreover, the high temperatures (up to 1900 K) allow higher thermal efficiencies (up to 50%).

    From HIGH FRONTIER by Philip Eklund


    MInature ReacTor EnginE. MITEE is actually a family of engines. These are small designs, suitable for launching on existing boosters. You can find more details here.

    Basic MITEE
    Exhaust Velocity9,810 m/s
    Specific Impulse1,000 s
    Thrust14,000 N
    Thrust Power68.7 MW
    Mass Flow1 kg/s
    Total Engine Mass200 kg
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power3 kg/MW

    The baseline design is a fairly conventional NTR. Unlike earlier designs it keeps the fuel elements in individual pressure tubes instead of a single pressure vessel, making it lighter and allowing slightly higher temperatures and a bit better exhaust velocity.

    Monatomic H
    Monatomic-H MITEE
    Exhaust Velocity12,750 m/s
    Specific Impulse1,300 s
    Thrust2,350 N
    Thrust Power15.0 MW
    Mass Flow0.18 kg/s
    Total Engine Mass200 kg
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power13 kg/MW

    This advanced design works at a lower chamber pressure so that some of the H2 propellant disassociates into monatomic hydrogen, although the chamber temperature is only slightly greater. The drawback is that this reduces the mass flow through the reactor, limiting reactor power.

    Exhaust Velocity17,660 m/s
    Specific Impulse1,800 s
    Thrust1,700 N
    Thrust Power15.0 MW
    Mass Flow0.10 kg/s
    Total Engine Mass10,000 kg
    Uranium 235
    ReactorSolid Core
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power666 kg/MW

    The use of individual pressure tubes in the reactor allows some fuel channels to be run at high pressure while others are run at a lower pressure, facilitating a hybrid electro-thermal design. In this design cold H2 is heated in the high pressure section of the reactor, is expanded through a turbine connected to a generator, then reheated in the low pressure section of the reactor before flowing to the nozzle. The electricity generated by the turbine is used to break down more H2 into monatomic hydrogen, increasing the exhaust velocity. Since this is a once through system there is no need for radiators so the weight penalty would not be excessive.

    Liquid Core

    Liquid Core 1
    Exhaust Velocity16,000 m/s
    Specific Impulse1,631 s
    Thrust7,000,000 N
    Thrust Power56.0 GW
    Mass Flow438 kg/s
    Total Engine Mass70,000 kg
    Uranium 235
    ReactorLiquid Core
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power1 kg/MW
    Liquid Core 2
    Exhaust velocity14,700 to 25,500 m/s

    Nuclear thermal rocket / liquid core fission. Similar to an NTR-GAS, but the fissionable core is merely molten, not gaseous. A dense high temperature fluid contains the fissionable material, and the hydrogen propellant is bubbled through to be heated. The propellant will be raised to a temperature somewhere between the melting and boiling point of the fluid. Candidates for the fluid include tungsten (boiling 6160K), osmium (boiling 5770K), rhenium (boiling 6170K), or tantalum (boiling 6370K).

    Liquid core nuclear thermal rockets have a nominal core temperature of 5,250 K (8,990°F).

    The reaction chamber is a cylinder which is spun to make the molten fluid adhere to the walls, the reaction mass in injected radially (cooling the walls of the chamber) to be heated and expelled out the exhaust nozzle.

    Starting up the engine for a thrust burn will be complicated and tricky, shutting it down even more so. Keeping the fissioning fluid contained in the chamber instead of escaping out the nozzle will also be a problem.


    Exhaust Velocity19,620 m/s
    Specific Impulse2,000 s
    Thrust20,000 N
    Thrust Power0.2 GW
    Mass Flow1 kg/s
    Total Engine Mass1,000 kg
    Uranium 235
    ReactorLiquid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power5 kg/MW
    Propulsion SystemLARS
    Exhaust Velocity10,300 m/s
    Specific Impulse1,050 s
    Thrust11,000,000 N
    Thrust Power56.6 GW
    Mass Flow1,068 kg/s
    Total Engine Mass9,000 kg
    Uranium 235
    Thrust DirectorNozzle
    Wet Mass226,000 kg
    Dry Mass45,000 kg
    Mass Ratio5.02 m/s
    ΔV16,623 m/s

    Nuclear thermal rocket / liquid annular reactor system. A type of NTR-LIQUID. You can find more details here

    The molten fissioning uranium is held in tubes which are spun to provide centifugal gravity. This keeps the uranium from escaping out the exhaust, mostly. Seeded hydrogen propellant is injected down the spin axis where it is heated by the nuclear reaction then escapse out the exhaust nozzle.

    These engines have a specific impulse ranging between 1,600 to 2,000 seconds, and an internal temperature between 3,000K and 5,000K

    Droplet Core

    Droplet Core
    Reactor inner diameter1 m
    Reactor outer diameter2 m
    Reactor inner length3 m
    Reactor outer length4 m
    Engine length13 m
    (no shadow shield)
    (with shadow shield)
    Engine mass
    (no shadow shield)
    6,800 kg
    Engine mass
    (with shadow shield)
    21,200 kg
    Engine pressure500 atm
    Internal temp6,000K
    Isp2,000 sec
    Exhaust velocity19,600 m/s
    Engine power1,500 MWth
    Thrust333,000 N

    The data is from Droplet Core Nuclear Rocket (1991).

    The main draw-back is that developing such an engine will be just as hard as developing a gas core nuclear thermal engine. But it has much lower performance. So why bother?

    This propulsion system straddles the line between liquid-core and vapor-core. Much like how vapor-core straddles the line between liquid-core and gas-core. Instead of the uranium fuel being in the form of gaseous vapor, it is instead in the form of a fog of droplets.

    Droplet core engines have a specific impulse between 1,500 and 3,000 seconds and an internal temperature between 5,000K and 7,000K. The specific impulse is enhanced because the nuclear energy is strong enough to dissociate some (20%) of the hydrogen molecules of propellant into atomic hydrogen. The propellant flow rate can be between 1 to 1,000 kilograms per second.

    The temperature depends upon the pressure inside the chamber. The design shown assumes a pressure of 500 atmospheres, where the melting point of uranium is 1,400K and the boiling point is 9,500K. This is enough to heat the hydrogen propellant to 6,000K and gives a specific impulse of 2,000 seconds.

    The chamber is about one meter in diameter and three meters tall.

    At the top molten uranium with a temperature of around 2,000K is injecting through the unfortunately named "atomizer." In this case the term has nothing to do with nuclear physics, but more to do with Victorian perfume spray bottles. The droplets are from five to ten microns in size, and enough are sprayed into to create a critical mass. The upper 1.5 meters of the chamber is clad in neutron reflectors, so about 70 to 80% of the power generated occurs here. The next meter has only partial neutron reflectors, and the lower half meter has no neutron reflectors at all. Naturally the neutron flux is highest in the part with the most reflectors.

    In the upper half of the chamber hydrogen propellant bleeds in from the walls, but in the lower half high pressure tangential jets spray a flood of hydrogen. Like vapor-core and open-cycle-gas-core the frantically fissioning uranium is intimately mixed with the hydrogen propellant. This gives an almost three orders of magnitude improvement on heat transfer area (i.e., about a thousand times better than a solid-core nuclear engine). The propellant is heated not only by heat radiation, but also by heat conduction of hydrogen gas in direct contact with the uranium drops. A whopping 30% to 40% of the fission energy is transferred to the propellant.

    The tangential spray in the lower half of the chamber does two things: [1] help keep the blasted uranium drops from splattering on the walls and [2] create a vortex that will assist capturing uranium so it can be re-used instead of losing it out the exhaust nozzle. That stuff is both deadly and expensive, you don't want any un-burnt uranium escaping. The report calculates that the uranium loss will be less than 50 kilograms per mission.

    About half a meter from the bottom of the chamber the tangential hydrogen jets are replaced with molten lithium-6 jets. The vortex makes the hot uranium drops hit the relatively cool lithium layer. This chills the uranium so the drops mix with the lithium. The mixture is captured at the bottom and sent to a fuel separator. The unburnt uranium is sent back to the top for another trip through the chamber while the lithium is sent back to the lithium jets.

    The engine has a very high thrust-to-weight ratio. A 1,500 MWth engine with 333,000 Newtons of thrust would have a T/W of 5.0. Though actually that drops to 1.6 once you add the radiation shadow shield so the crew doesn't die. If my slide rule is not lying to me, this means the described engine has a mass of 6.8 metric tons with no radiation shield, and a mass of 21.2 metric tons with (or a shield mass of 14.4 metric tons).

    This particular engine would have about 20 kg of uranium in the reaction chamber at any given time, and 100 kg total fuel. As mentioned before the report predicts it will lose about 50 kg out the exhaust nozzle over an entire mission.

    Vapor Core

    Vapor Core
    Thrust Power1.6 GW
    Exhaust velocity9,800 to
    11,800 m/s
    Thrust330,000 n
    Propellant mass flow30 kg/sec
    Reactor thermal power1,400 to
    1,800 MW
    Total engine mass6.83 tonne
    Fuel element mass total1.35 tonne
    Forward reflector mass0.60 tonne
    Aft reflector mass0.51 tonne
    Radial reflector mass2.47 tonne
    Radiation shield mass0.9 tonne
    Total reactor mass5.83 tonne
    Misc. engine
    component mass
    0.9 tonne
    Uranium Hexafluoride
    ReactorVapor Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power4 kg/MW

    This is sort of an intermediate step in learning how to design a full-blown Gas Core Nuclear Thermal Rocket. It is basically a solid core NTR where the solid nuclear fuel elements are replaced by chambers filled with uranium235 tetrafluoride vapor. The engine is admirably compact with a nicely low critical mass, and an impressive thrust-to-weight ratio of 5-to-1. However the specific impulse / exhaust velocity is only slightly better than a solid core.

    In other words, the system is not to be developed because it has fantastic performance, but because it will be an educational step to building a system that does.

    The specific impulse is around 1,280 seconds and the internal temperature is between 6,000K and 8,000K.

    The uranium fuel is kept physically separate from the hydrogen propellant, so the exhaust is not radioactive.

    A 330,000 newton thrust NVTR would have a core with almost 4,000 fuel elements, with a core radius of 120 cm, core height of 150 cm, and 1,800 MW. Criticality can be achieved with smaller cores: a core volume five times smaller with radius of 60 cm, height of 120 cm, and power of 360 MW.

    Data is from Conceptual Design of a Vapor Core Reactor Rocket Engine for Space Propulsion by E.T. Dugan, N.J. Diaz, S.A. Kuras, S.P. Keshavmurthy, and I. Maya (1996).

    ForwardBeryllium oxide15 cm0.60 tonne
    AftC-C Composite25 cm0.51 tonne
    RadialBeryllium oxide15 cm2.47 tonne
    CORE: 2000 fuel elements
    Radius0.5 m
    Height1.5 m
    Fuel channel per element12 to 32
    Hydrogen channel per element12 to 32
    Critical mass20 kg
    Hydrogen pressure100 atm
    UF4 pressure100 atm
    Fuel center temperature4,500 K
    Design Values
    Pump Flowrate (Total)75.20 lbm/s
    Pump Discharge Pressure3,924 psia
    Pump Efficiency80.01%
    Turbopump RPM70,000 RPM
    Turbopump Power (each)9,836 HP
    Turbine Inlet Temperature481 deg-R
    Turbine Pressure Ratio1.69
    Turbine Flow Rate (each)33.77 lbm/s
    Reactor Thermal Power1,769 MW
    Fuel Element and Reflector Power1,716 MW
    Nozzle Chamber Temperature5,580 deg-R
    Chamber Pressure (Nozzle Stagnation)1,500 psia
    Nozzle Expansion Area Ratio500:1
    Vacuum Specific Impulse (Delivered)997.8 sec
    Heat Loads
    Nozzle-con (total)30.05 MW
    Nozzle-div (total)22.97 MW
    Reflector (total)35.0 MW
    Typical NVTR Engine Parameters
    Nozzle Area Ratio500
    Fuel Pressure100 atm
    Average Fuel Temperature4000 K
    Maximum Element Heat Flux420 W/cm2
    Nomial Element Length150 cm
    Fuel Volume Fraction0.15
    Coolant Volume Fraction0.15
    Moderator Volume Fraction0.70
    Fuel Element Power0.9 MWt
    Element Heat Transfer Area2141 cm2
    Reactor Core L/D1.5
    Fuel Channel Diameter0.142 cm
    Fuel Channel Sectional Area0.0158 cm2
    Total Fuel Channel Area Per Element0.505 cm2
    Fuel Element Sectional Area3.464 cm2
    Element Diameter (across flats)2.2 cm
    Coolant Channel Diameter0.142 cm
    Coolant Channel Sectional Area0.0158 cm2
    Total Coolant Channel Area Per Element0.505 cm2
    Core Volume1.2 m3
    Core Volume Density1,500 MW/m3
    Fuel Element Mass, Total1.35 MT
    Forward Reflector Mass0.60 MT
    Aft Reflector Mass0.51 MT
    Radial Reflector Mass2.47 MT
    Radiation Shield Mass0.90 MT
    Total Reactor Mass5.83 MT
    Misc. Engine Components Mass0.9 MT
    Total Engine Mass6.83 MT
    Engine F/W5.0

    Gas Core

    Remember, all nuclear thermal rockets are using nuclear energy to heat hydrogen propellant for rocket exhaust. The hotter the reactor core, the more the propellant is heated, and the higher the specific impulse and exhaust velocity. That means the rocket has more delta-V go travel to more distant places, and also can carry more payload.

    The problem is that the reactor is made out of matter, and above a certain temperature the reactor melts. Go higher and the reactor vaporizes into gas. Solid-core nuclear thermal rockets keep the temperature below the melting point, which means they top out at a specific impulse of 1,200 seconds or so. Admittedly this is better than the pathetic 450 seconds you can squeeze out of a conventional chemical rocket. But it is still not high enough to really open up the exploration of the solar system.

    If you allow the uranium to reach a temperature where it melts you can get up to a specific impulse of 2,000 seconds or so. This is a liquid-core nuclear thermal rocket. You spin the reaction chamber around the thrust axis to make the hot bubbling liquid uranium stick to the chamber walls instead of escaping out the exhaust.

    But if you want to crank it up to the max you have to let the uranium reach temperatures where it vaporizes into white-hot gas. This can get up to a whopping 3,500 seconds of specific impulse.

    The drawback is trying to keep all that expensive and deadly gas from shooting out the exhaust bell. Which isn't easy.

    Closed-Cycle gas-core NTR try to have it both ways. They enclose the nuclear fury of gaseous uranium in solid quartz-crystal containers to keep the exhaust non-radioactive. Which is counter-productive since the whole idea was to let everything vaporize for maximum heat output. The end result is the specific impulse will be about half of what it could be.

    Open-cycle gas-core NTR just let it all hang out. Radioactive fission-products vapor escapes out the exhaust making it very unhealthy to be anywhere near the rocket when it is thrusting. But it has the maximum specific impulse. Since that enriched uranium is hideously expensive you want to at least make a cursory effort to keep it in the reaction chamber as long as possible. You do not want un-burnt uranium escaping, you want it all burnt in the reaction chamber. The general rule is that as long as the hydrogen mass expelled is 25 to 50 times a high as the uranium mass expelled the uranium losses are within acceptable limits. The various open-cycle designs use different strategies to make the hydrogen to uranium exhaust flow ratio as high as possible.

    Closed Cycle

    Gaseous Core NTR closed 1
    Exhaust Velocity20,405 m/s
    Specific Impulse2,080 s
    Thrust445,000 N
    Thrust Power4.5 GW
    Mass Flow22 kg/s
    Total Engine Mass56,800 kg
    Uranium Hexafluoride
    ReactorGas Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power13 kg/MW
    Gaseous Core NTR closed 2
    Thrust Power0.6 to 231 GW
    Exhaust velocity10,800 to 31,400 m/s
    Thrust117,700 to 14,700,000 n
    Engine mass30 to 300 tonne
    Engine T/W0.4 to 5.0
    Operating Pressure400 to 1600 atm
    NASA report nuclear lightbulb
    Thrust Power3.7 GW
    Engine Power4.6 GW
    Exhaust velocity18,300 m/s
    Thrust409,000 n
    Engine mass32 tonne
    Engine T/W1.3
    Operating Pressure500 atm
    Propellant mass flow22.3 kg/s
    Liberty Ship
    Propulsion SystemNuclear Lightbulb
    Exhaust Velocity30,000 m/s
    Specific Impulse3,058 s
    Thrust/Engine5,340,000 N
    Number Thrustersx7
    Thrust37,380,000 N
    Thrust Power560.7 GW
    Mass Flow1,246 kg/s
    Total Engine Mass378,000 kg
    Uranium Hexafluoride
    Wet Mass2,700,000 kg
    Dry Mass1,600,000 kg
    Mass Ratio1.69 m/s
    ΔV15,697 m/s
    Specific Power0.67 kg/MW

    Closed-cycle gaseous core fission / nuclear thermal rocket AKA "Nuclear Lightbulb". Similar to an open-cycle gas core fission rocket, but the uranium plasma is confined in a fused quartz chamber. It is sort of like a child's classic Easy-Bake Oven. Except that there is propellant instead of cake mix and the light bulb is full of fissioning uranium instead of electricity.

    You can read more about this on the Unwanted Blog in the posts here, here, and here.

    The good news is that unlike the open-cycle GCNR it does not spray glowing radioactive death there is no uranium escaping in the exhaust. The bad news is that the maximum exhaust velocity is halved, as is the Delta-V. Yes, I did ask some experts if it was possible to make some kind of hybrid that could "shift gears" between closed and open cycle. Sorry, there would be no savings over just having two separate engines.

    The maximum exhaust velocity is halved because you are trying to have it both ways at once. The higher the propellant's temperature, the higher the exhaust velocity and rocket Delta-V. Unfortunately solid core nuclear reactors have a distressing habit of vaporizing at high temperatures (as do all material objects). Gas core reactors are attempting to do an end run around this problem, by having the reactor start out as high temperature vapor. But adding the quartz chamber is re-introducing solid material components into the engine, which kind of defeats the purpose. The only thing keeping this from utter failure is the fact that quartz does not heat up as much due to the fact it is transparent.

    Even with the restrictions, it seems possible to make a closed-cycle GCNR with a thrust to weight ratio higher than one. This would allow using the awesome might of the atom to boost truely massive amounts of payload into Terra orbit, without creating a radioactive wasteland with every launch. See the GCNR Liberty Ship for an example. The Liberty Ship can boost in one launch more payload than any given Space Shuttle does in the shuttles entire 10 year operating life. Then the Liberty Ship can land and do it again.

    The ideal solution would be to somehow constrain the uranium by something non-material, such as a mangnetohydrodynamic force field or something like that. Alas, currently such fields can only withstand pressures on the order of the breeze from a flapping mosquito, not the 500 atmospheres of pressure found here. But since researchers are working along the same lines in their attempts to make a fusion reactor, this may change. And I did find a brief reference to something called an "MHD choke" in reference to slowing the escape of uranium into the exhaust stream of an open-cycle gas core rocket. I'm still trying to find more information on that.

    The high pressure is to ensure the uranium vapor is dense enough to sustain a fission reaction.


    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.

    VCR light bulb fission
    VCR light bulb fission
    Exhaust Velocity19,620 m/s
    Specific Impulse2,000 s
    Thrust56,400 N
    Thrust Power0.6 GW
    Mass Flow3 kg/s
    Total Engine Mass72,566 kg
    Uranium Hexafluoride
    ReactorGas Core
    RemassSeeded Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power131 kg/MW

    Most fission reactors avoid meltdown, but the vapor core reactor (VCR) runs so hot (25000 K) that its core vaporizes.

    At this temperature, the vast majority of the electromagnetic emissions are in the hard ultraviolet range. A “bulb” transparent to this radiation, made of internally-cooled a-silica, bottles the gaseous uranium hexafluoride, while letting the fission energy shine through.

    The operating pressure is 1000 atm. The UF6 fuel is prevented from condensing on the cooled wall by a vortex flow field created by the tangential injection of a neon “buffer” gas near the inside of the transparent wall.

    In a generator mode, the UV uses photovoltaics to generate electricity. In a propulsion mode, the UV heats seeded hydrogen propellant, which exits at a specific impulse of 2000 seconds.

    From High Frontier by Philip Eklund
    NASA Report

    The pictured engine is the "reference engine" described in the report Studies of Specific Nuclear Light Bulb and Open-Cycle Vortex-Stabilized Gaseous Nuclear Rocket Engines. I will give the high-lights from the report, but for all the boring nitty-gritty details you'll have to read the report yourself. Please note that as always, "down" is the direction the thrust is going (to the right in the blueprint) and "up" is the opposite direction (to the left). I apologize for the use of imperial units instead of metric. The report is in imperial, and it is too much of a pain to change everything.

    The important statistics: Specific impulse is 1870 seconds (18,300 m/s), thrust 409,000 newtons, engine mass 32,000 kg, thrust-to-weight ratio 1.3.

    The other stats. The total volume of the reaction chambers (cavities) is 170 cubic feet. There are seven cavities, each six feet long. The cavity pressure is 500 atmospheres. The specific impulse is 1870 seconds (report says can theoretically be from 1500 to 3000 seconds). The total propellant flow (including seed and nozzle transpirant coolant flow) is 49.3 pounds per second. The thrust is 92,000 pounds. The engine power is 4600 megawatts. The engine weight is 70,000 pounds. If one can design a variable-throat-area nozzle (instead of fixed-area) this will result in a major decrease in the required chamber pressure during startup.


    The basic configuration is seven separate unit cavities surrounded by moderator-reflector material in between each cavity (beryllium oxide) and surrounding the entire cavity array (graphite). Each cavity is 6.0 feet long and the total volume of all seven cavities is 169.8 cubic feet. The cavity pressure is 500 atmospheres due to criticality and fuel density considerations.


    In each lightbulb, a critical mass of gaseous uranium creates thermal radiation. The thermal radiation can pass through the transparent quartz crystal walls of the lightbulb, but the uranium vapor cannot. This means no lethal uranium enters the exhaust. Hydrogen propellant flowing over the lightbulb is heated to high temperatures by the thermal radiation and is expelled out the rocket nozzles, producing thrust. The hydrogen is "seeded" with tungsten dust because it too is ordinarily transparent to thermal radiation. The seeding makes it opaque, and allows it to be heated. Seven "lightbulbs" are used instead of one, since that increases the total lightbulb radiating area by about 2.2 times.

    Transparent quartz walls

    The transparent quartz wall of the lightbulb contains lots of coolant channels. This is because the quartz is mostly transparent to thermal radiation, but not totally. And fissioning uranium produces an awful lot of thermal radiation. I told you that nuclear lightbulb designers were trying to have it both ways. The coolant channels are marked "circumferential coolant tubes" in the diagram below.

    Inside a lightbulb

    Inside the lightbulb, neon buffer gas is used to create a vortex ring to suspend the gaseous nuclear fuel (a "radial inflow" vortex). The vortex ring looks like an elongated donut (I know it looks like two separate blobs above, that's due to the fact the diagram is a cross-section). One of the important jobs done by the neon buffer gas is to prevent the 42,000°R uranium plasma from making contact with the lightbulb walls. This would be very bad, as the walls would be instantly vaporized. The neon passes along the lightbulb walls, bends round the end caps, then travels down the long axis of the lightbulb (right down the center of the vortex ring). When it reaches the fore end cap, it is removed from the lightbulb through a port (marked "thru-flow" in diagram above).

    The removed neon is very hot, and contains unburnt uranium and fission products. It is cooled by mixing with low-temperature neon, which condenses the unburnt uranium vapor into hot liquid uranium. The liquid uranium is separated from the neon by a centrifuge and sent back into the vortex (at point marked "fuel injection"). The neon is cooled further then it too is sent back into the vortex (at point marked "buffer gas injection"). While examining the blueprint, I noticed that the centrifuges, and indeed the entire uranium fuel delivery system is conspicuous by its absence. Probably classified.

    Note that the centrifuges is a neat solution to the problem of fission fragments clogging up the fuel. In essence, this design has its own built-in nuclear fuel reprocessing plant. Of course the nasty fission fragments will have to be stored and eventually disposed of.

    Lightbulb dimensions

    The total volume inside all the lightbulbs is 84.9 cubic feet, which is 12.1 cubic feet per lightbulb. The radius of the uranium fuel containing region is 85% of the radius of the transparent wall. While the fissioning uranium fuel has a core temperature of 42,000° Rankine, the outer surface is only at 15,000° Rankine.

    Propellant flow in a lightbulb

    The propellant is assumed to exit with a temperature of 80% of the fuel temperature, or 12,000° Rankine. This is because the quartz transparent walls will reflect about 15% of the thermal radiation back inside. By some compilcated reasoning that you will find in the report, the total thermal radiation from the lightbulbs is 4.37 x 106 BTU/sec. The hydrogen propellant has an "enthalpy" of 1.033 x 105 BTU/pound at 12,000°R. So by dividing the two, one discovers that the entire engine can support a propellant flow rate of 42.3 pounds per second, which means 6.07 lb/sec for each of the seven cavities.

    If that last paragraph confused you, let me explain. As a simple example, if a pound of hydrogen at 5°R contains 2 BTUs ("enthalpy"), and the engine puts out 6 BTU per second, then obviously the engine can heat up 6 / 2 = 3 pounds of hydrogen per second. Why do we care? If you multiply the propellant flow rate by the exhaust velocity you will discover the engine's thrust value. And that's a number we do care about.

    The tungsten dust that the propellant is seeded with has a particle diameter of 0.05 microns. The seed density is 1.32 x 10-2 lb/ft3, which is about 3.9 percent of the inlet propellant density. This can probably be reduced if tungsten dust was in the form of thin flat plates instead of spherical particles.

    The hydrogen propellant enters the pressure shells from the fore end (see "Primary Circuit Inlet" in pressure shell diagram below). A bit is bled off from small H2 flow ports in order to pressurize the interior of the shells, circulating to provide coolant to the engines and machinery. But most of it is fed into the turbopump, then injected into the cavities. Since the fore end of each cavity is almost blocked off by the butt end of the lightbulb, there is only a narrow rim to inject the hydrogen.

    In the diagram to the right, you can see how the propellant is fed from the pink pipe into the pink-and-gold wedge-shaped injectors. I presume there are three injectors per cavity, spraying into the clear area between the transparent wall's coolant manifolds and buffer gas injectors.

    Uranium fuel

    The total fissioning uranium in all seven vortexes be about 25.2 pounds of uranium (about 3.6 pounds per cavity). You would ordinarily need more to ensure nuclear criticality, but the required amount is brought down by the beryllium oxide neutron reflector encasing each cavity. The average uranium fuel density is 0.409 lb/ft3. The total density of the neon-uranium mix inside the vortex is about 0.56 lb/ft3. A unit of neon gas will spend about 3.8 seconds inside the cavity. A unit of uranium will spend about 19 seconds inside the cavity. This implies a uranium fuel flow rate of 0.19 lb/sec per cavity.

    According to my slide rule, if the array of seven cavities is producing 4,600 megawatts, it means that the array is burning a miniscule total of 0.055 grams (0.00012 pounds) of uranium fuel per second (0.0079 grams per cavity per second). It still needs the full 3.6 pounds per cavity to be present in order to burn the fraction of a gram.

    The theoretical maximum specific impulse possible is 2230 seconds. Due to this designs incomplete expansion, transpiration coolant flow in the nozzle, presence of tungsten seeding, and friction losses the specific impulse is reduced to 84% or 1870 seconds. Total propellant flow (allowing for tungsten seeds and transpiration cooling) is 49.3 lb/sec. This would result in a thrust of 92,000 pounds force. For complicated reasons you can find in the report, this implies that the exhaust nozzles are 0.0875 feet in diameter at the throat expanding to 2.04 feet diameter at the exit.

    Uranium refueling

    Careful readers may have noticed how the description avoids mentioning the details on how one gets the uranium into the lightbulbs. This is because it is quite a difficult problem, and each of the proposed solutions has drawbacks. The basic problem is old reliable: all the atomic fireworks inherent in 235U will happen if you merely let too much of it accumulate in one place. You have to store it diffuse and somehow bring it together in the lightbulb.

    Method #1 Store it as uranium hexafluoride gas. This would be in large tanks of low pressure (i.e., low density) and with the tanks full of neutron absorbing foam. Pump enough into the lightbulb, a chain reaction will start, and well before the reaction reaches 13,000°R the uranium will have separated from the fluorine.

    The problem is that now you have the insanely dangerous task of dealing with 13,000°R fluorine gas. At room temperature the blasted stuff will violently react with any element in the known universe except helium and neon. A temperature of 13,000°R makes it about 13,000 times as deadly. It will explosively corrode away anything solid in its path like molten lead on facial tissue. Chemist Derek Lowe sarcastically notes that "At seven hundred freaking degrees, fluorine starts to dissociate into monatomic radicals, thereby losing its gentle and forgiving nature." You can read more about the suicidal risk of dealing with hot fluorine in his amusing blog post.

    Method #2 Store it as sub-critical chunks of uranium, melt them, and inject the molten uranium into the lightbulb. Uranium melts at 1403°K, which is difficult but not impossible. The plan is to somehow turn the molten uranium into a sort of aerosol mist suspended in hot neon.

    The problem is that the molten uranium wants to plate itself all over the melter and the aerosol spray equipment. Which is annoying if the material in question is something like lead, but disasterous if the material is radioactive and fissionable.

    Method #3 is to store the uranium cold as finely divided dust. As dust it is pumpable, injectable, and it will not plate over everything. Inside the lightbulb the uranium dust will be rapidly heated to vaporization by the nuclear reaction. This method does not have any major problems, except for the common problem of how to protect the transparent wall from being vaporized by the heat.

    Again, the uranium delivery system seems to be totally missing from the blueprint. The only bit present is the short stub of the injector at the top of each lightbulb.

    Pressure shells

    The entire engine is encased in two nested pressure shells constructed of filament-wound fiberglass. The inside of the inner shell is pressurized to 500 atmospheres. Hydrogen propellant enters through a 0.5 foot diameter duct at the fore end (aka "Primary Circuit Inlet"). There are seven 0.4 foot diameter holes in the aft end for the engine nozzles, one at zero degrees off-axis, the other six at 60°. The pressure shell can be separated into two parts along the flange at the point of maximum diameter, to allow an engineer or waldo manipulator access to the engine interior. This point is also where the rear structural grid protrudes from the interior, this is where the engine bolts onto the structural frame of the spacecraft to transmit the engine thrust.

    If you look at the large blueprint, you will see that parts of the rear structural grid penetrate the cavities to support the end-caps of the quartz lightbulbs.

    Coolant system

    The plumbing for the coolant system is rather complicated (translation: I don't understand it all). Click for larger image. You can use this diagram along with the large blueprint to attempt to puzzle out what all the pipes are for. Basically the propellant enters the system through the "Primary circuit inlet" (at lower left of plumbing diagram, and in the pressure shell diagram above) and leaves the system via the "Propellant injection" arrow, where the propellant is heated by the lightbulbs in the cavity and jets out the exhaust nozzles. In between, the propellant frantically threads its way over every single other engine component in a desperate attempt to cool them off.

    Cross sections

    Here are a set of cross sections through the cavities. The one on the left is zoomed in on the cavity interior, the other two gradually zoom out.

    UAC Report

    The information comes from a series United Aircraft Corporation reports written mostly by Thomas L. Latham. There are more reports than the ones I've used.

    The reference design had seven cells with six surrounding the center cell. The entire engine was sized to fit into the Space Shuttle cargo bay. It was also sized at 4.6 gigawatts, 409,000 Newtons, and a specific impulse of 1,860 seconds in order to avoid the need for external heat radiators. At this level no radiators are required for the moderator or pressure vessel, open-cycle cooling will suffice. Above a specific impulsle of 1,860 seconds radiators will be needed or the engine will melt.

    If the specific impulse is above 2,500 seconds the nozzle throats will require their own cooling system.

    The hydrogen propellant is seeded with tiny tungsten particles due to the unfortunate fact that hydrogen is transparent to the frequencies emitted by the nuclear reaction. Otherwise the chamber walls would be heated instead of the propellant, which is the exact opposite of what we want. The fissioning U235 or U233 fuel also emits ultraviolet light that degrades the transparency of the enclosing quartz "lightbulb." The researchers were experimenting with seeding the uranium with something that would turn the UV into infrared in order to protect the quartz. Happily the ionizing radiation does expose the degraded quartz to a radiation damage annealing effect that restores transparency to some extent.

    The fuel is in the form of Uranium Hexafloride.

    The average dose rate in the filament-wound fiberglas pressure vessel was calculated to be 0.17 mrad/sec. This would allow about six full-power runs of 1000-sec duration (about 17 minutes) before the total dose became 1000 mrad, the estimated allowable dosage before degradation of the laminate strength commences.


         This is basically the propellant, passing from the propellant tanks to be heated by the nuclear light bulbs, and then rushing through the exhaust nozzles to provide thrust. Along the way it provides some cooling for various items.
         Starting at the tank, the primary hydrogen pump sends it through a H2-H2 heat exchanger for preheating (and providing additional heat rejection for the Secondary Hydrogen Circuit). It passes through a H2-Ne heat exchanger to cool off the neon gas in the Neon And Fuel Circuit. It passes through the Fuel And Neon Separator. A turbine then sends it through the Solid Moderator and End-Wall Liners. Somewhere along the line it is seeded with tungsten microparticles so the hydrogen will be heated by the nuclear light bulbs.
         Finally it experiences extreme Direct Heating from the nuclear light bulbs, and exits through the exhaust nozzles.
         Basically the coolant system. It runs cooling hydrogen over the pressure vessel, nozzles, flow divider, tie rods, liner tubes, and the transparent walls of the quartz light bulbs (during shutdown it also cools the Fuel And Neon Separator).
         The now-hot hydrogen passes through a H2-H2 heat exchanger to give the heat to the space radiator. The lukewarm hydrogen passes through a second H2-H2 heat exchanger to cool down further and preheat the propellant hydrogen.
         The Neon Make-Up supply keeps the neon pressure in the circuit at the required level. The uranium-235 Fuel Make-Up keeps the amount of fuel droplets in the circuit at the required level. Both are fed into the Fuel Cavity in the interior of the quartz light bulbs to create the furious nuclear reaction (unless engine shut-down is in progress, then the Fuel Control Valve closes to shut off the uranium). The reaction provides the direct heating to the Primary Hydrogen Propellant Circuit. Some of the neon goes through the Cavity Bypass Flow.
         Only a fraction of the uranium undergoes fission. So the neon/uranium that comes out of the Fuel Cavity is sent through the Neon and Fuel Separator to strain the uranium out of the neon gas. The neon is cooled which makes the uranium gas condense into liquid droplets. The two are separated by a centrifuge. The neon is cooled further by the H2-Ne heat exchanger.
         The neon goes to the Neon Pump, the uranium goes to the Fuel Pump and the cycle begins anew.

    For additional details see Ref. 5 (Nuclear studies of the nuclear light bulb rocket engine).

    Neon supply is the Neon Make-Up supply, keeping the neon pressure in the circuit at the required level. It is fed into the Fuel Cavity (Unit Cavity) tangentally just inside the quartz light bulb Transparent Wall. This creates the neon-uranium vortex.

    The Fuel distillation canister is the Fuel Make-Up. It is fed by the Fuel Pump into the fuel injection duct, introducing it into the Fuel Cavity (Unit Cavity). This creates the furious nuclear reaction inside the quartz light bulb, providing the direct heating to the Primary Hydrogen Propellant Circuit.

    The mixture of hot neon, unburnt gaseous uranium fuel, and fission products exits the Fuel Cavity via the Exhaust Duct (about two meters long). Not shown is how cool neon is introduced into the entire length of the exhaust duct to [1] cool the exhaust from 6550 K to 1500 K, [2] prevent the exhaust from severely damaging the exhaust duct, [3] condense the gaseous uranium into liquid uranium droplets, and [4] ensuring that the uranium droplets condense inside the neon gas, instead of on the walls of the exhaust duct causing a nuclear reaction.

    The 1500 K neon-uranium droplet flow is sent to the Neon and Fuel Separator (Separator) where the two are isolated by a centrifuge. The neon is cooled by the H2-Ne heat exchanger and goes to the Neon Pump. The uranium fuel goes to the Fuel Pump. Alternatively the uranium is distilled to separate out the silicon seeding and the uranium is deposited in the fuel distillation canister.

    Values for weight flow rates, temperature, and volume flow rates are indicated at various stations in the system.

    In the Neon and Fuel Separator, the seven exhaust duct inlet pipes from the seven nuclear light bulb unit cavities enter from the left. They enter two inlet plenums: four inlet pipes on the top plenum and three on the bottom. Each plenum has an injection slot delivering the gas mix into the separator cavity, with a velocity of 500 m/s at a steep angle designed to spin the gas. The spin centrifugally separates the uranium from the neon, at about 100,000 g's. The uranium is harvested by uranium collector tubes on the separator wall, while the neon is harvested by an outlet pipe on the separator's long axis. The separator cavity and uranium collector tubes have to be maintained at or above 1,500 K, or the uranium will condense on them. This will not only clog the thing up, but if enough uranium plates out it will accumulate a critical mass with regrettable results.


    Startup Sequence

    1. Fill hydrogen ducts and neon system from storage to a pressure equal to approximately 20 atmospheres
    2. turn on neon recirculation pump
    3. inject fuel until critical mass is reached
    4. increase power level and adjust flow rates and cavity pressure to maintain criticality and limit component temperatures to tolerable level
    5. inject propellant seeds when 10 percent of full power is reached
    6. increase power to desired operating leve

    The paper looks at two "power ramps", going from cold to full power in 60 seconds or a more leisurely 600 seconds. Below a temperature of 15,000°R the fusing uranium is heating up the hydrogen propellant mainly by convection. Above 15,000°R the uranium heats the propellant by infrared thermal radiation.

    Since convection does such a pathetic job of transfering heat, most of the fission energy goes to heating up the uranium dust instead of the propellant. In about five seconds flat the uranium reaches 12,000°R, and vaporizes from dust into red-hot gas. Then at 15,000°R thermal radiation takes over and the uranium temperature rises more slowly (which you can see by the way the curve starts flattening out). At 60 or 600 seconds (depending upon which power ramp you used) the uranium is at the nominal temperature of 45,000°R. It won't rise any higher unless the engine is exploding or something rude like that.

    As previously mentioned the hydrogen propellant is pretty much transparent to thermal radiation, which is most unhelpful. Normally the infrared will shoot right through the hydrogen without heating it up. So tungsten dust is seeded into the propellant to soak up the thermal radiation and heat the propellant by conduction. Any thermal radiation that misses the seeding will hit the far wall of the propellant chamber, which is also the beryllium oxide moderator (BeO) helping to keep the uranium fissioning. The thermal heating of the BeO is nothing but wasted energy but the seeding is doing the best it can. The BeO is designed so it can handle up to 2,400°R.

    Since the BeO moderator outweighs the uranium dust by several orders of magnitudue, it takes far longer to heat up. As you can see from the graph the uranium fuel starts heating up after only 0.03 seconds but the BeO doesn't even start heating until 10 seconds, about 300 times longer. The uranium gets up to nominal temperature in 60 seconds but the BeO takes 300 seconds. And the BeO only gets up to 2,400°R while the uranium is smokin' at 45,000°R. That is for the 60 second ramp. The 600 second ramp has both the uranium and BeO all warmed up at the same time, only because 600 seconds gives the BeO time to catch up.

    However, the shorter 60 second ramp is desireable, because the 600 second ramp wastes precious propellant. Take the propellant mass required for a standard 20 minute burn at full power. The 60 sec ramp requires an additional 2.7% propellant as startup wastage. The 600 sec ramp requires a whopping 27% additional, which is totally unacceptable. What, do I look like I am made of propellant? The paper says it might be possible to reduce the ramp time down to 6 seconds, in the interest of reducing the propellant startup wastage even further (presumably to 0.27%).

    The critical mass of uranium-235 fuel required in the quartz tubes increases during the ramp up. It requires 18.6 pounds at zero power up to 30.9 lbs at full power. For the 60 second ramp up full power initially happens at 28.2 lbs, but rises to 30.9 lbs at 300 seconds. This is because at 60 seconds the BeO moderator has only warmed up about two-thirds of the way to its max temperature. Apparently once the BeO is fully warmed up the critical mass rises.

    When the paper was read, one of the attendees was skeptical about pressure. Specifically if the pressure of the uranium/neon mix is not the exact same as the pressure of the hydrogen propellant, the pressure differential will shatter the quartz tube like dropping an old-school incandescent lightbulb on a concrete floor. The paper authors insisted that the two pressures could be balanced rapidly enough to prevent that unhappy state of affairs. They say that a differential of two or three atmospheres will shatter the blasted tube, so they want to keep the diff under 2/3rds atm. Yikes, I didn't know that! That would instantly ruin the propulsion system, and spray everybody and everything close by with fissioning uranium.



    Shutdown Sequence

    1. Close Fuel Injection Control Valve (turn off the uranium)
    2. Begin Linear Decrease in Propellant Flow Rate (propellant flow past light bulbs to exhaust nozzles)
    3. Begin Linear Increase in Radiator Flow Rate (flow from coolant heat exchanger to radiator)
    4. Maintain Secondary Circuit (flow of hydrogen coolant) and Cavity Neon Flow (buffer gas flow inside the quartz light bulbs) at Full Power Value.

    Once the engine shut-down sequence is initiated, it takes six seconds for the power level to drop to zero. It only takes 0.8 seconds for power level to drop to 0.01 of full power, during which time the contained uranium fuel drops from the steady-state level of 13.65 kg down to 11.5 kg.

    Pulsed Close Cycle

    This is from Pulsed Plasma-Core Rocket Reactors (from Research on Uranium Plasmas and their Technological Applications page 52) (1970)

    This is actually quite clever. Dr. Winterberg was trying to address the two main problems with open-cycle gas core reactors: preventing unburnt U235 from escaping out the exhaust nozzle, and dealing with wear and tear on the engine from the horrifically high operating temperatures. His solution was to pulse the reaction.

    Now remember nuclear fission 101: when a thermal neutron crashes into a uranium 235 nucleus, the nucleus is split into fission fragments, and nuclear energy is released. Oh, and it also emits several neutrons, which keep the chain reaction going.

    You want to burn as much of the U235 as possible, that stuff's expensive. If you can't burn it all in the rocket chamber, the next best thing is to try and catch the unburnt U235 before it escapes out the exhaust and re-use it.

    where ΔNu/Nu is the percent of U235 that was successfully burnt in a fission reaction, σf is the fission neutron cross section, φ is the neutron flux, and τ is the fuel confinement time (or lifetime in the reactor if engine is a solid core NTR).

    In other words, improving the amount of U235 burnt means increasing the the amount of uranium atoms getting in the way of neutrons, increasing the number of neutrons for the uranium to get in the way of, and increasing how long the uranium atoms are playing demolition derby with the neutrons. Which is kind of obvious if you think abou it.

    So if a bog-standard nuclear power reactor had a neutron cross section σf of 10-22 cm3, a neutron flux φ of 1015, and was turned on τ of 103 seconds (16.6 minutes); then the ΔNu/Nu u235 burnup would be 0.0001 or 0.01%.

    Open-cycle gas core nuclear rockets are really bad at confining the fuel for any reasonable length of time. τ is really low. To make up for this you have to increase the neutron cross section or the neutron flux. Or both. Increasing the neutron cross section means drastically increasing the chamber pressure to make the U235 cloud more dense, which means more mass for a heavy-duty pressure chamber, which sends the engine's thrust-to-weight ratio gurgling down the toilet. Increasing the neutron flux means more neutron heating of the engine, or even enough neutron heating to actually vaporize the engine.

    Dr. Winterberg noted that while increasing the neutron cross section is probably out of the question, there might be a way to manage an increase in neutron flux. The neutron heating of the engine relies upon duration. The longer the engine is exposed to the neutron flux, the hotter it gets. Reduce the exposure time and you reduce the engine temperature rise. In other words: Pulse the reaction. You can use a fantastically high neutron flux as long as the duration of the flux is short enough so that the engine does not overheat. Wait for the engine to cool off then you can pulse again.

    Taken to extremes, you'll have the equivalent of an Orion drive, where the reaction is less a slow energy release and more like an Earth-Shattering Kaboom. A bomb in other words. But Dr. Winterberg saw there was a lot of performance improvement possible using pulses much less violent than bomb-level. More to the point, improvements that would allow the engine to get away with having very short confinment times.

    The engine will use a high neutron flux to pulse a series of "soft" nuclear detonations. This will have the following advantages:

    • The high neutron flux will increase the U235 burnup rate to the point where you can get away with a shorter required fuel confinement time. The short pulse will ensure that the neutron flux does not vaporize the engine.
    • The higher temperatures created in the reaction chamber will increase the exhaust velocity and specific impulse something wonderful. Again the short pulse will prevent the temperatures from damaging the engine
    • Pulse operation allows starting the chain reaction from a uranium-propellant mixture at high density with a small critical mass. This allows the reaction chamber and the rest of the engine to be smaller than other gas-core designs.
    • Using pulsed operation allows using a dynamic system to separate the fuel from the propellant, meaning to prevent uranium from escaping out the exhaust so the engine will be more closed-cycle than open cycle. Details to follow.

    Figure (a)

    The reactor vessel is surrounded by a conventional nuclear reactor (not shown). It is designed to be powered up then powered down rapidly, to create an intense pulse of neutron flux (something like φ = 1014/cm2 sec) inside the reactor vessel.

    Valve V opens, and into the reactor vessel is injected a slug of hydrogen propellant, containing a subcritical piece of U235. The valve snaps shut behind the slug.

    Note that the U235 is off-center inside the slug, further away from the exhaust nozzle than most of the propellant. This is so when the U235 explodes, most of the hydrogen propellant will be blown out the exhaust nozzle before any of the fissioning U235 reaches the nozzle exit.

    The U235 is still fully embedded in the propellant, none of the U235 is exposed. This is so all the nuclear explosion energy hits the propellant, instead of frying the interior of the reactor vessel.

    A tiny "trigger" piece of U235 is injected at high velocity down pipe T.

    Figure (b)

    When the trigger enters the subcritical U235, the surrounding nuclear reactor simultaneously pulses an intense neutron flux. The assembly becomes prompt critical and a small nuclear explosion ensues. This heats the hydrogen propellant which is pushed out the exhaust nozzle, creating thrust. Propellant on the other sides of the explosion protecting the reactor vessel from thermal radiation.

    Figure (c)

    Just before the unburnt U235 and fission fragment cloud escapes through the exhaust nozzle, the nozzle plug P closes the nozzle. The nozzle plug has to close at a rate of about 104 cm/sec, which can be done with a plug driven by pressurized gas. The hot propellant / unburnt U235 / fission fragment cloud is trapped inside the reactor vessel. This is sucked out of the reactor vessel through pipe E.

    The gases are sent through a heat radiator to be cooled off. Then they enter a fuel-propellant reprocessing plant. This separates the three ingredients. The fission fragments are disposed of. The hydrogen propellant is sent to the propellant tank. The unburnt U235 is carefully fabricated into subcritial fuel masses, being very careful not to let a critical mass accidentally accumulate. The subcritical masses are sent to the fuel storage unit.

    Open Cycle

    The open-cycle gas core engine has a radioactive exhaust, there is no getting around it. So the first thing you have to do is estimate the radiation hazard and ensure the crew has adequate radiation shielding.

    The second thing to do is find a design that does not wastefully allow expensive un-burnt uranium to escape out the tailpipe. Again the general rule is that as long as the hydrogen mass expelled is 25 to 50 times a high as the uranium mass expelled the uranium losses are within acceptable limits. The various open-cycle designs use different strategies to make the hydrogen to uranium exhaust flow ratio as high as possible.

    Crew radiation dose from plume of Gas-Core rocket

    In the open-cycle gas-core nuclear rocket concept the heat source is fissioning uranium gas. This released heat is radiated to and absorbed by the hydrogen propellant, The heated propellant is exhausted through a nozzle, producing thrust. The fission fragments that are formed and the unfissioned uranium fuel are also exhausted into the vacuum of space. As the plume is formed, the crew is exposed to gamma radiation from the fission fragments in the plume.

    The radiation dose to the crew from the fission fragments in the plume can be separated into two components. Component one results from the fact that there is a microscopic amount of plume material that has sufficient kinetic energy to flow back towards the vehicle. Some of this material will strike and stick to the vehicle. Since this material will contain fission fragments, these gamma radiation sources will stay with the crew throughout the entire trip and this dose could represent a significant source of radiation. Masser(3) has estimated this dose and has concluded it would be less than 10-3 rem for a typical manned Mars mission.

    Component two of the dose results from the fission fragment distribution throughout the entire plume volume and is potentially much larger than component one. Since the plume contains over 99 percent of the exhausted material, 99 percent of the fission fragments will be in the plume. It is the purpose of this paper to estimate the radiation dose rate and total dose to the crew from the fission fragments in the plume for four specific missions to the planet Mars.

    Another source of radiation is caused by the delayed decay of the fission fragments that are passing through the nozzle. This includes delayed neutrons which can cause secondary fissioning and gamma's. This source, however, has not been included. There is another radiation source associated with the gas-core reactor, that of the reactor core. This radiation source, along with solar radiation, must be ultimately considered when total dose rates to the crew are evaluated. This study, however, is concerned only with that part of the total radiation problem that arises from the fission fragments in the plume volume.

    3. Masser, C. C., "Radiation Hazzard from Backflow of Fission Fragments from the Plume of a Gas-Core Nuclear Rocket," Research on Uranium Plasmas and Their Technological Applications, SP- 236, 1971, NASA, Washington, D.C.

    (ed note: the equations used to draw these graphs are in the document. I didn't bother to include them since they involve calculus. The radiation doses in the graphs give spacecraft designers the radiation shielding requirements)

    1. For the most probable fission fragment retention, time of 100 seconds, and crew nozzle separation of 100 meters, the radiation dose varied from 170. to 36. rem for the 80 and 200 day round trip times respectively. Five centimeters of lead shielding would reduce the radiation dose by two orders of magnitude, thereby protecting the crew. The increase in vehicle weight would be insignificant. For example, a shield of five centimeters thickness and four meters in diameter would add 7120 kilograms to the vehicle gross weight of 0.94 million kilograms. Also additional attenuation is available In the form of liquid hydrogen propellant, spacecraft structure, nuclear fuel, equipment, and stores.

    2. For the trip times included in this analysis the total radiation dose to the crew is proportional to the energy required for the mission. Therefore, within the ranges used in this analysis one can estimate the crew radiation dose by knowing the energy needed for the mission.

    3. For the crew-nozzle separation of 100 meters, approximately 50 percent of the plume radiation is received from the first 0.1 kilometer into the plume. This percentage is increased to 90 percent for 1 kilometer and 100 percent for 100 kilometers into the plume.

    4. For an 80 day round trip to Mars, with a crew-nozzle separation distance of 100 meters, the radiation dose varied from about 0.5 to 1670. rem for fission fragment retention times of 10,000 and 10 seconds, respectively.

    5. For all cases, increasing the crew distance from 100 to 200 meters from the nozzle exit reduced the unshielded radiation dose by half.

    General Open Cycle
    Open Cycle
    Propulsion SystemGas Core NTR
    Exhaust Velocity35,000 m/s
    Specific Impulse3,568 s
    Thrust3,500,000 N
    Thrust Power61.2 GW
    Mass Flow100 kg/s
    Total Engine Mass200,000 kg
    Uranium Hexafluoride
    ReactorGas Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power3 kg/MW
    Engine mass30-200 tonne
    T/W 11.9 to 1.8
    Open Cycle 2
    Propulsion SystemGas Core NTR
    Exhaust Velocity50,000 m/s
    Specific Impulse5,097 s
    Thrust5,000,000 N
    Thrust Power0.1 TW
    Mass Flow100 kg/s
    Uranium Hexafluoride
    RemassLiquid Hydrogen
    Specific Power2 kg/MW
    Engine mass30-200 tonne
    T/W 17.0 to 2.5
    Open Cycle 3
    Thrust Power GW
    Exhaust velocity25,000 to 69,000 m/s
    Thrust19,600 to 108,000 n
    Engine mass40 to 110 tonne
    T/W0.05 to 0.10
    Operating Pressure400 to 2000 atm
    Open Cycle MAX
    Exhaust Velocity98,000 m/s
    Specific Impulse9,990 s
    Thrust3,000,000 N
    Thrust Power0.15 TW
    Mass Flow31 kg/s
    Total Engine Mass15,000 kg
    Uranium Hexafluoride
    Propulsion SystemGas Core NTR
    Exhaust Velocity35,316 m/s
    Specific Impulse3,600 s
    Thrust3,500,000 N
    Thrust Power61.8 GW
    Mass Flow99 kg/s
    Uranium Hexafluoride
    RemassLiquid Hydrogen
    Wet Mass433,000 kg
    Dry Mass268,000 kg
    Mass Ratio1.62 m/s
    ΔV16,943 m/s

    Gaseous core fission / nuclear thermal rocket AKA consumable nuclear rocket, plasma core, fizzer, cavity reactor rocket. The limit on NTR-SOLID exhaust velocities is the melting point of the reactor. Some engineer who obviously likes thinking "outside of the box" tried to make a liability into a virtue. They asked the question "what if the reactor was already molten?"

    Gaseous uranium is injected into the reaction chamber until there is enough to start a furious chain reaction. Hydrogen is then injected from the chamber walls into the center of this nuclear inferno where it flash heats and shoots out the exhaust nozzle.

    The trouble is the uranium shoots out the exhaust as well. This not only makes the exhaust plume dangerously radioactive but it also wastefully allows expensive unburnt uranium to escape before it contributes to the thrust.

    The reaction is maintained in a vortex tailored to minimize loss of uranium out the nozzle. Fuel is uranium hexaflouride (U235F6), propellant is hydrogen. However, in one of the designs, U235 is injected by gradually inserting into the fireball a long rod of solid uranium. The loss of uranium in the exhaust reduces efficiency and angers environmentalists.

    In some designs the reaction chamber is spun like a centrifuge. This encourages the heavier uranium to stay in the chamber instead of leaking into the exhaust. This makes for a rather spectacular failure mode if the centrifuge's bearings seize.

    You can find more details here.

    The thermal radiation from the fission plasma is intended to heat the propellant. Alas, most such engines use hydrogen as the propellant, which is more or less totally transparent to thermal radiation. So the thermal stuff goes sailing right through the hydrogen (heating it not at all) then striking the reaction chamber walls (vaporizing them).

    To remedy this sorry state of affairs, gas-core designers add equipment to "seed" the propellant with something opaque to thermal radiation. Most of the reports suggest tungsten dust, with the dust size about the same as particles of smoke, about 5% to 10% seeding material by weight. The seeding absorbs all but 0.5% of the thermal radiation, then heats up the hydrogen propellant by conduction. The chamber walls have to cope with the 0.5%.

    Most of the reports I've read estimate that the reaction chamber can withstand waste heat up to 100 megawatts per square meter before the chamber is destroyed. For most designs this puts an upper limit on the specific impulse at around 3,000 seconds.

    However, if you add a heat radiator to cool the reaction chamber walls and the moderator surrounding the reaction chamber, you can handle up to about 7,000 seconds of specific impulse. The drawback is the required heat radiator adds lots of mass to the engine. A typical figure is of the total mass of a gas core engine with radiator, about 65% of the mass is the radiator.

    Another fly in the ointment is that the proposed seeding materials turn transparent and worthless at about the 10,000 second Isp level. To push the specific impulse higher a more robust seeding material will have to be discovered. Since current heat radiators cannot handle Isp above 7,000 seconds, robust seeding is not a priority until better radiators become available.

    Yet another challenge is that 7% to 10% of the fission plasma power output is not in the form of thermal radiation, but instead neutrons and gamma rays. Which the propellant will not stop at all, seeded or not. This will penetrate deep into the chamber walls and moderator (since gamma-rays are far more penetrating than x-rays), creating internal waste heat.

    Sub 3,000 Isp designs deal with radiation heat with more regenerative cooling. Higher Isp need even more heat radiators.

    Most designs in the reports I've read use 98% enriched uranium-235 (weapons-grade). The size of the reaction chamber can be reduced somewhat by using uranium-233 according to this report.

    The reaction chamber size can be reduced by a whopping 70% if you switch to Americium-241 fuel according to this report. The drawback is the blasted stuff is $1,500 USD per gram (which makes every gram that escapes un-burnt out the exhaust financial agony). The short half-life means there is no primordial Americium ore, you have to manufacture it in a reactor via nuclear transmutation. The report estimates that for a 6 month brachistochrone trajectory the spacecraft would need about 2,000 kilograms of the stuff. Which would be a cool three million dollars US. I'm sure the price would drop if dedicated manufacturing sites were established to create it.

    If used for lift off it can result in a dramatic decrease in the property values around the spaceport, if not the entire country. An exhaust plume containing radioactive uranium is harmless in space (except to the crew) but catastrophic in Earth's atmosphere.

    Amusingly enough, this is the best match for the propulsion system used in the TOM CORBETT: SPACE CADET books. However the books are sufficiently vague that it is possible the Polaris used a nuclear lightbulb. According to technical advisor Willy Ley, "reactant" is the hydrogen propellant, but the books imply that reactant is the liquid uranium.



    The temperature limitations imposed on the solid core thermal rocket designs by the need to avoid material melting can be overcome, in principle, by allowing the nuclear fuel to exist in a high temperature (10,000 — 100,000 K), partially ionized plasma state. In this so-called "gaseous- or plasma-core" concept, an incandescent cylinder or sphere of fissioning uranium plasma functions as the fuel element. Nuclear heat released within the plasma and dissipated as thermal radiation from its surface is absorbed by a surrounding envelope of seeded hydrogen propellant that is then expanded through a nozzle to provide thrust. Propellant seeding (with small amounts of graphite or tungsten powder) is necessary to insure that the thermal radiation is absorbed predominantly by the hydrogen and not by the cavity walls that surround the plasma. With the gas core rocket (GCR) concept Isp values ranging from 1500 to 7000 s appear to be feasible [Ref. 26]. Of the various ideas proposed for a gas core engine, two concepts have emerged that have considerable promise: an open cycle configuration, where the uranium plasma is in direct contact with the hydrogen propellant, and a closed-cycle approach, known as the "nuclear light bulb engine" concept, which isolates the plasma from the propellant by means of a transparent, cooled solid barrier.

    Porous Wall Gas Core Engine

    The "open cycle," or "porous wall," gas core rocket is illustrated in Fig. 9. It is basically spherical in shape and consists of three solid regions: an outer pressure vessel, a neutron reflector/moderator region and an inner porous liner. Beryllium oxide (BeO) is selected for the moderator material because of its high operating temperature and its compatibility with hydrogen. The open cycle GCR requires a relatively high pressure plasma (500 — 2000 atm; 1 atm = 1.013 × 105 N/m2 ) to achieve a critical mass. At these pressures the gaseous fuel is also dense enough for the fission fragment stopping distance to be comparable to or smaller than the dimensions of the fuel volume contained within the reactor cavity. Hydrogen propellant, after being ducted through the outer reactor shell, is injected through the porous wall with a flow distribution that creates a relatively stagnant non-recirculating central fuel region in the cavity. A small amount of fissionable fuel (1/4 to 1 % by mass of the hydrogen flow rate) is exhausted, however, along with the heated propellant.

    Because the uranium plasma and hot hydrogen are essentially transparent to the high energy gamma rays and neutrons produced during the fission process, the energy content of this radiation (~7—10% of the total reactor power) is deposited principally in the solid regions of the reactor shell. It is the ability to remove this energy, either with an external space radiator or regeneratively using the hydrogen propellant, that determines the maximum power output and achievable Isp for the GCR engines. To illustrate this point, an open cycle engine with a thrust rating of 220 kN (50,000 lbf) is considered. We assume that 7% of reaction energy Prx reaches the solid, temperature-limited portion of the engine and that the remainder is converted to jet power at an isentropic nozzle expansion efficiency of ηj. Based on the realtionships between Isp, reactor power, and propellant flow rate (ṁp) given below.

    (ed note: elsewhere in this website, ṁ is called "m-dot")

    0.93·Prx(MW) = 4.9×10-6·F(N)·Isp(s) / ηj

    0.93·Prx(MW) = 4.9×10-5·ṁp(kg/s)·Isp2(s) / ηj

    a 5000 s engine generating 7500 MW of reactor power will require a flow rate of 4.5 kg/s at rated thrust. If the hydrogen is brought into the cavity at a maximum overall operating temperature of 1400 K, no more than 1.2% of the total reactor power (~17% of the neutron and gamma power deposited in the reactor structure) can be removed regeneratively (ṁp cp ΔT ≈ 90 MW). Total removal requires either (1) operating the sold portions of the engine at unrealistically high temperatures (>11,000 K at ṁp = 4.5 kg/s) or (2) increasing the propellant flow rate substantially to 36.8 kg/s (at 1400 K), which reduces the engine's Isp to 1750 s. "Closed cooling cycle" space radiator systems have been proposed [Ref. 27] as a means of maintaining the GCR's operational flexibility. With such a system, adequate engine cooling is possible even during high Isp operation when the hydrogen flow is reduced. Calculations performed by NASA/Lewis Research Center [Ref. 28] indicate that specific impulses ranging from 3000 to 7000 s could be attained in radiator-cooled, porous wall gas core engines.

    The performance and engine characteristics for a 5000 s class of open cycle GCRs are summarized in Table 4 for a range of thrust levels. The diameter of the reactor cavity and the thickness of the external reflector/moderator region are fixed at 2.44 m and 0.46 m, respectively, which represents a near-optimum engine configuration. The engine weight (Mw) is composed primarily of the pressure vessel (Mpv); radiator (Mrad); and moderator (Mmod).

    Table 4
    Characteristics of 5000 s Porous Wall Gas Core Rocket Engines







    1. For a hydrogen cavity inlet temperature of 1400 K and a heat deposition rate that is 7% of the reactor power, the ratio of radiated to total reactor power is a constant equal to 5.8%.
    2. The weight of the spherical pressure vessel is based on a strength-to-density value of 1.7×l05 N-m/kg [Ref. 29] which Is characteristic of high strength steels.
    3. Used in these estimates is a radiator specific mass of 145 kg/MW [Ref. 28] which is based on a heat rejection temperature of 1225 K and a radiator weight per unit surface area of 19 kg/m2
    4. Density of BeO is 2.96 mT/m3.

    By fixing the engine geometry in Table 4 the mass of the BeO moderator remains constant at 36 mT. However, the pressure vessel and radiator weights are both affected by the thrust level. While the radiator weight increases in proportion to the extra power that must be dissipated at higher thrust, the reason for the increase in pressure vessel weight is slightly more subtle. For a constant Isp engine an increase in thrust is achieved by increasing both the reactor power and hydrogen flow rate. In order to radiatively transfer this higher power to the propellant, the uranium fuel temperature increases, necessitating an increase in reactor pressure to maintain a constant critical mass in the engine. Accommodating this increased pressure leads to a heavier pressure vessel. (In going from 22 kN to 440 kN, the engine pressure rises from 570 atm to 1780 atm).

    As Table 4 illustrates, the moderator is the major weight component at lower thrust levels (<110 kN) while the radiator becomes increasingly more important at higher thrust. At thrust levels of 220 kN and above, the radiator accounts for more than 50% of the total engine weight. There is therefore a strong incentive to develop high temperature (~1500 K) liquid metal heat pipe radiators that could provide significant weight reductions in the higher thrust engines.

    Table 4 also shows an impressive range of specific powers (alphas) and engine thrust-to-weight ratios for the thrust levels examined. The F/Mw ratio for the 22 kN engine is over two orders of magnitude higher than the 5000 s nuclear-powered MPD electric propulsion system proposed in the Pegasus study [Ref. 30]. For manned Mars missions the higher acceleration levels possible with the GCR can lead to significant (factor of 5) reductions in trip time compared to the Pegasus system.


    (ed note: calculated estimates of gas core nuclear rocket engine weights for specific impulses ranging from 3000 to 7000 seconds and for engine thrusts ranging from 4400 to 440,000 newtons. Contains useful equations for calculating the mass of various engine components.)


    Virtually all existing or proposed rocket propulsion engines can be categorized as either high-thrust systems or high-specific-impulse systems. What is really needed for fast interplanetary travel is both characteristics, namely a high specific impulse (3000 sec or greater), and an engine thrust-weight ratio that is in the range from to 10-1. The characteristics of a gas-core nuclear rocket engine are examined in this study to see how closely it meets these requirements.

    Calculations were carried out to estimate gas-core engine weights for specific impulses ranging from 3000 to 7000 seconds and thrust levels from 4.4×103 to 4.4×105 newtons. A vapor-fin space radiator operating at 1100 K was incorporated into the engine system to dispose of waste heat not regeneratively removed by the hydrogen propellant. The total engine weight was composed of the individual weights of the radiator, the reactor moderator-reflector materials, the pressure shell, the nozzle, and the propellant turbopump. The study produced the following results and conclusions:

    1. Gas-core engines have the potential of producing a specific mass in the range 0.6 to 0.02 kilogram of weight per kilowatt of thrust power.

    2. For a specific impulse of 5000 seconds and a thrust of 1.1×105 newtons, engine weight is estimated to be 91 000 kilograms. This weight is composed of about equal proportions of radiator, moderator, and pressure shell weights. For the entire range of specific impulses and thrust levels of this study, engine weight varied from 35 000 to 380 000 kilograms.

    3. Engine weight increases with increasing specific impulse and with increasing thrust level. For a given specific impulse, higher thrust-weight ratios occur at higher thrust levels because engine weight does not increase as fast as the thrust does.

    Figure 1(a) illustrates schematically how this basic notion might be translated into a rocket engine. It is not unreasonable to picture this kind of engine as a nuclear "sun" with the central fireball and surrounding gas flow contained within a chamber' surrounded by structural materials. The analogy is not exact, of course, because the heat generation is due to nuclear fission rather than fusion. However, in both cases tha amount of energy that can be generated in, and released from, the fireball is essentially unlimited. There is, however, a limitation on how much energy can be absorbed by the hydrogen and turned into thrust without overheating the cavity wall or the exhaust nozzle. It is the amount of energy that reaches various solid, temperature-limited regions of the engines that ultimately limits the power generation and therefore the specific impulse.

    The proposed reactor shown in figure 1(a) is basically spherical. It is composed of an outer pressure vessel, a region of heavy-water reflector, a high-temperature beryllium moderator region, an inner heavy-water moderator, and finally a porous or slotted cavity liner. Approximately 7 to 10 percent of the reactor power is deposited in these solid regions of the reactor due to attenuation of high-energy gamma and neutron radiation. This heat is removed either by a coolant in an external space radiator loop, or regeneratively by the hydrogen propellant before it enters the central reactor cavity, The beryllium region is operated at a temperature of about 1300 K and the radiator at 1100 K.

    Uranium metal would have to be injected into this high-pressure region. Once inside the cavity, the uranium vaporizes and rises to temperatures sufficient to thermally radiate the energy that is generated by the fissioning uranium. A possible fuel injection technique might consist of pushing a thin rod of solid uranium metal at a high velocity through a shielded pipe (perhaps made of cadmium oxide) that penetrates the moderator. Some cooling of the uranium fuel and the shielded passage may be required to remove the heat that would be generated in the fuel as it passes through the moderator region. A 100-kilogram force would be required to drive a 0.15-centimeter diameter wire into a cavity with a pressure of 5.07×107 newtons per square meter. As it enters the cavity, the uranium instantly vaporizes and rises in temperature to about 55 000 K. Reactor startup could be achieved by first establishing the hydrogen flow. Next uranium particles would be blown into the dead cavity region to achieve nuclear criticality. The power would then be increased to a level sufficient to vaporize the incoming uranium rod.

    The seeded hydrogen is heated solely by absorbing the thermal radiation from the fissioning uranium fireball. The cavity walls receive only about 1 or 0.5 percent of the thermal radiation from the fireball. This wall protection is accomplished by introducing about 1 percent by weight of a seeding material such as graphite or tungsten particles into the hydrogen. This same technique is used in the nozzle region to reduce the hydrogen radiation heat load and the hydrogen temperature near the nozzle wall to tolerable levels. Seed concentrations of about 1 to 10 percent are required here. Figure 1 shows that some cold hydrogen can be introduced through the nozzle walls directly from the plenum at the downstream end of the engine if it is required. This would tend to reduce the specific impulse.


    The specific impulse of a gas-core rocket engine is limited by the fraction of the reactor power that reaches the solid, temperature-limited portions of the engine, and by how that heat is removed. It is an unavoidable characteristic of the nuclear fission process that about 7 to 10 percent of the energy release is high-energy gamma and neutron radiation that will go through the hydrogen gas but be stopped in the surrounding solid reactor structure.

    This energy that is deposited in the moderator can be regeneratively removed by the incoming hydrogen propellant. There is, however, a limit to how much heat the hydrogen can accommodate. For a 3000-second specific impulse engine, 7 percent of the reactor power will heat all the hydrogen propellant to 2800 K before it enters the reactor cavity. To achieve a higher specific impulse would require the solid parts of the engine to operate at an unrealistically high temperature. If the reactor materials, including the porous cavity wall, were limited to a little over 1000 K and if only regenerative cooling were used, the specific impulse would be limited to 2000 seconds.

    Higher specific impulses are possible by using an external radiator to reject part of the moderator heat to space. The radiator is shown schematically in figure 1. To bring the hydrogen into the reactor cavity at 1000 K for a specific impulse of 5000 seconds would require that the hydrogen remove no more than about 1 percent of the reactor power from the moderator, as shown in figure 2. The remaining 6 to 9 percent would have to be removed by the radiator loop.

    It appears that the ultimate limitation on specific impulse of a gas-core engine will depend on the ability to absorb the thermal radiation from the fuel in the hydrogen so that the cavity wall and the nozzle wall do not receive an excessive heat flux. Based on current estimates of the optical absorption and emission properties of the gases involved, a recent Lewis in-house study indicates that the maximum specific impulse is in the range 5000 to 7000 seconds.


    The engine weight analysis used for this study is the same as was presented in reference 4, except for the addition of a space radiator and the elimination of a specific equation for fuel volume as a function of the hydrogen-to-uranium-mass-flow ratio (how many units of hydrogen propellant are expended in the exhaust before one unit of uranium is lost). The engine weight is taken to be the sum of the individual weights of the moderator, pump, nozzle, pressure shell, and radiator:


    An initial series of calculations were made to select a "best" cavity diameter and moderator thickness combination. This preliminary optimization was done at values of specific impulse (5000 sec) and thrust (4.4×104 N) that are centered in the ranges covered in this study. One cavity diameter and one moderator thickness were selected on this basis, and then held constant for all subsequent variations of specific impulse and thrust. Thus, after this initial reactor optimization, the moderator weight was not a variable in this study.

    Engine Pressure

    In order to calculate the weights of the nozzle, turbopump, and pressure shell, it was necessary to calculate the pressure required to have a critical mass in the engine. This was obtained from the following equation:


    where P is the reactor pressure in atmospheres, Mc, is the critical mass in kilograms, F is the engine thrust in newtons, Isp is the specific impulse in seconds, Dc is the reactor cavity diameter in meters, and VF is the fraction of the reactor cavity filled with fuel. Equation (2) is more general than the form used in reference 4 where a specific relation between fuel volume fraction and hydrogen-touranium-mass-flow ratio was used to eliminate VF from equation (2). The present study was carried out for a fuel volume fraction of 0.25. Recent fluid mechanics experiments using air/air indicate that this value should be attainable for hydrogen-to-uranium-flow ratios in the range 100 to 400.

    Nozzle, Turbopump, And Pressure Shell


    where the component weights are in kilograms, F is thrust in newtons, Isp is specific impulse in seconds, P is reactor pressure from equation (2) in atmospheres, and Rs is the inside radius of the pressure shell in meters.

    The radiator weight estimate was based on a recent study of a vapor-fin for space power systems. The vapor-fin design would weigh 290 kilograms per megawatt of radiated power, based on operating the radiator at 945 K. For this study it was assumed that the same radiator, or at least one of the same weight per unit surface area (19 kg/m2 of plan form area), could be operated at 1100 K. This gives a weight of 145 kilograms per megawatt of radiated power:


    Equations (2) to (6) were used to obtain the weight of each engine component. Equation (1) was used to obtain the total engine weight. For this study, calculations were carried out for specific impulses of 3000, 5000, and 7000 seconds, and for engine thrusts from 4.4×103 to 4.4×105 newtons.

    It may be necessary to operate the radiator at a pressure less than that of the reactor cavity in order to keep the lightweight vapor-fin design. For example, the pressure stress in the radiator tube walls would range from 10.14×107 to 50.7×107 newtons per square meter for internal tube pressures ranging from 10.14×106 to 5.07×107 newtons per square meter, respectively. This same pressure stress range could be reduced by a factor of 3 by increasing the tube wall thickness such that the overall radiator weight would increase by about 20 percent. In an actual engine design, one might not choose to do this, but instead operate the radiator at a lower pressure than that of the reactor. This would then require a pump to increase the radiator discharge pressure to that inside the reactor pressure vessel.


    The engine weight results are presented and discussed in this section. First, the effect of varying the cavity diameter and the moderator thickness is presented. Based on these results, one cavity diameter and one moderator thickness are selected for the remainder of the calculations. For this fixed engine geometry, the effect of thrust level on engine weight is determined for a specific impulse of 5000 seconds. Next, the effect of specific impulse on engine weight is presented over a range of thrust levels. Finally, these results are presented in terms of a parameter commonly used to describe lowthrust propulsion devices, engine specific mass, which is the ratio of engine weight to thrust power (in kg/kW).

    Effect of Cavity Diameter and Moderator Thickness

    Changes in cavity diameter or in moderator thickness cause two effects on engine weight. One effect is that the weight of moderator material is changed. The other effect is that the uranium density required for criticality is changed. This changes the required reactor pressure, which, in turn, results in a change in the pressure shell weight.

    These two influences on engine weight tend to oppose each other. For example, reducing the moderator thickness reduces the moderator weight, but increases the pressure required for criticality. Thus, there is some optimum moderator thickness that gives a minimum engine weight. It is possible, however, that the engine pressure at this minimum-weight geometry would be unrealistically high, so that one might choose to operate at some near but off-optimum configuration that has a somewhat lower pressure.

    Engine weight was calculated for five combinations of cavity diameter and moderator thickness. The results are shown in figure 3. The critical mass requirements are listed in table I. These engine weight calculations were carried out for a specific impulse of 5000 seconds and a thrust level of 4.4×104 newtons. Both of these values are centered within the ranges covered in this study.

    Cavity diameters of 2.4, 3.6, and 4.9-meters were used with a constant moderator thickness of 0.76 meter. Moderator thicknesses of 0.61, 0.76, and 0.91 meter were used with a constant cavity diameter of 3.6 meters. Within these ranges, reductions in either parameter caused a decrease in engine weight but an increase in engine pressure. A cavity diameter of 2.4 meters with a moderator thickness of 0.76 meter produced the lightest engine, which weighed 64 000 kilograms. The reactor pressure for this engine was 7.8×107 newtons per square meter.

    Further reduction of cavity diameter below 2.4 meters would probably have produced a slightly lighter engine, but at the expense of an extremely high pressure. This is shown in figure 4. On the basis of these results, a 2.4-meter cavity diameter and a 0.76-meter moderator thickness were selected as representing a near-optimum engine configuration. The remaining calculations were carried out using this one engine geometry.

    Effects of Thrust Level on Engine Weight

    Higher thrust requires a heavier engine. The component weights are shown in figure 5 for engine thrust varying from 0 to 1.1×105 newtons at a specific impulse of 5000 seconds. For a thrust below about 5×104 newtons, the radiator weight is not too important, compared to the moderator and the pressure shell weights. At a thrust of 1.1×105 newtons, the radiator, pressure shell, and moderator each contribute about one-third of the total engine weight.

    For higher thrusts, the radiator weight begins to dominate. This is shown in table II. At a thrust of 2.2×104 newtons, the radiator only contributes 6400 kilograms to the total engine weight of 51 000 kilograms, or about 12 percent. At a thrust of 2.2×105 newtons, the radiator accounts for 64 000 kilograms out of 133 000 kilograms, or almost 50 percent. This indicates that for thrusts above 2.2×105 newtons, at this specific impulse of 5000 seconds, significant weight reductions can be achieved if higher temperature radiators can be developed. For example, the radiator weight could be cut in half by operating at 1300 K instead of 1100 K. All the calculations of this study were done for a radiator temperature of 1100 K.

    Effect of Specific Impulse on Engine Weight

    Higher specific impulses require heavier engines, at a given thrust level. This is shown in figure 6. For a thrust of 4.4×104 newtons, engine weights of 50 000, 64 000, and 73 000 kilograms are required for specific impulses of 3000, 5000, and 7000 seconds, respectively.

    At a specific impulse of 3000 seconds, a radiator may not be necessary. If the hydrogen propellant enters the reactor at 2800 K, it can regeneratively remove all the gamma and neutron heat deposition from the moderator region. This produces a lighter engine, as shown by the dashed curve in figure 6. Whether one would actually choose to operate the moderator at a little over 2800 K in order to achieve the lower weight would depend on a number of factors, such as the particular mission involved and the effect of moderator temperature on engine reliability and life. The solid curves in figure 6 are based on a hydrogen cavity inlet temperature of 1400 K. Table III lists the percent of reactor power that must be radiated away for this temperature.

    Gas-Core Specific Mass

    For low-acceleration systems such as electric thrusters, it is useful to characterize the propulsion device in terms of its specific mass. This parameter α is in kilograms of powerplant weight per kilowatt of thrust power. It can be related to the engine thrust-weight ratio as follows. The thrust power, or jet power as it is sometimes called, is given by 1/2 (F×Isp×g), which is simply the kinetic energy in the jet exhaust. SP Using this relation, the specific mass α, in kilograms per kilowatt, is


    where the specific impulse is in seconds and the engine thrust-weight ratio is dimensionless.

    Figure 7 shows the results of the present study presented on this basis. The specific mass of a gas-core engine varies from a high of 0.6 to a little less than 0.02 for specific impulses from 3000 to 7000 seconds and thrust levels from 4.4×10+3 to 4.4×10+5 newtons. Higher specific impulse or higher thrust produces a lower, and therefore better, specific mass.


    An analysis has been carried out to determine the characteristics of a low-thrust, high-specific-impulse, gas-core, nuclear rocket engine. The latest information on reactor critical mass requirements, radiant-heat-transfer properties, and fluid mechanics were used. For specific impulses above 3000 seconds, it was necessary to incorporate a space radiator as an engine system component. Engine weight was calculated for specific impulses ranging from 3000 to 7000 seconds, and for thrust levels from 4.4×103 to 4.4×105 newtons. Radiator weight estimates were based on an operating temperature of 1100 K. The calculations indicate the following results:

    1. Gas-core engines have the potential of producing a specific mass in the range 0.6 to 0.02 kilogram of weight per kilowatt of thrust power.

    2. For a specific impulse of 5000 seconds and a thrust of 1.1×105 newtons, engine weight is estimated to be 91 000 kilograms. This weight is composed of about equal proportions of radiator, moderator, and pressure shell weights. For the entire range of specific impulses and thrust levels of this study, engine weight varied from 35 000 to 380 000 kilograms.

    3. Engine weight increases with increasing specific impulse and with increasing thrust level. For a given specific impulse, higher thrust-weight ratios occur at higher thrust levels because engine weight does not increase as fast as the thrust does.


    This is from Mini Gas-Core Propulsion Concept by R. E. Hyland (1971) and A Study Of The Potential Performance And Feasibility Of A Hybrid-Fuel Open Cycle Gas Core Nuclear Thermal Rocket by Lucas Beveridge (2016).

    As previously mentioned open-cycle gas core engines solve the "reactor got so hot it vaporized" problem by starting out with the reactor already vaporized. The primary problem is how to prevent the uranium gas from prematurely escaping out the exhaust nozzle, but the secondary problems are pretty bad as well.

    To start the uranium fissioning, you need a certain amount of uranium at a certain density surrounded by enough neutron reflectors to kick stray neutrons back into play. Sadly, by definition, gaseous uranium has a much lower density than solid uranium. As it works out, for the engine to require a non-outrageous critical mass of uranium and a non-outrageous reaction chamber volume, the core pressure has to be very very high. Which will require a massive pressure vessel. Which makes the engine mass skyrocket. Which savagely cuts into the available payload mass and seriously degrades the engine's thrust-to-weight ratio.

    Oh, calamity and woe! How can this problem be remedied?

    Isp1,600 sec
    Exhaust Velocity15,696 m/s
    Thrust450 N
    Reactor Mass2,200 kg
    Pressure Shell Mass3,100 kg
    Radiator Mass4,930 kg
    Total Mass10,230 kg
    Outer Diameter1.22 m
    Cavity Diameter0.61 m
    233U Plasma Diameter0.43 m
    233U Plasma Mass1.42 kg
    233U Plasma Power
    4.5 MW
    233U Driver Power
    15.9 MW
    Total Engine Power
    20.4 MW
    Radiator Alpha310 kg/MW
    Pressure51 MPa
    Exhaust Temp4,000 to
    5,000 K


    Robert Hyland pondered the problem until the question arose "is it really necessary for all the uranium to be gaseous?"

    Hyland's solution was to embed a small solid core reactor in the walls of the chamber, as sort of a reactor layer. This is called the "driver core." It is far enough from the furious heat raging inside the chamber so it wouldn't melt. The driver core produces heat, but the important part is it produces neutrons. This makes the interior of the chamber so neutron-rich that the gaseous uranium does not have to be under such high pressure. In other words, the extra neutrons from the driver core lower the required critical mass of uranium gas inside the chamber.

    This allows the engine to get away with using a much less massive pressure vessel, which lowers the engine mass, which reduces the payload reduction and increases the thrust-to-weight ratio.

    Hyland said "I shall call him 'Mini-Gas Core.'" Lucas Beveridge called it the hybrid-fuel engine, since it uses both solid and gaseous uranium.

    Hyland scaled this to have an engine power of 20.4 MW, which implied a meager thrust of only 450 N. He thought it might be useful for unmanned space probes.

    So part of the total engine power is produced by the driver core (233U Driver Power or Psolid) and part of the total engine power is produced by the uranium plasma inside the chamber (233U Plasma Power or Pgas). Only Pgas is used to heat up the propellant to create thrust. Most of Psolid is just waste heat, a fraction of it is used to create neutrons to supercharge the uranium plasma. So heat radiators will be needed to get rid of the 15.9 megawatts worth of Psolid waste heat.

    Ptotal = Pgas + Psolid = 20.4 MW

    εgas = Pgas / Ptotal = 0.221

    εsolid = Psolid / Ptotal = 0.780

    εgas is the ratio of power in the 233U Plasma to the total. Hyland's Mini-Gas Core has a εgas of 0.221, or only 1/5th of the power is in the plasma. Beveridge found that was too low, and was focusing on a Low-ε engine with εgas = 0.51 and a High-ε engine with εgas = 0.673.


    Engine Common
    Thrust300,000 N
    Total Engine
    Power (Ptotal)
    3 GW
    Engine Mass<36,000 kg
    Low-ε Engine
    Isp1,600 sec
    Exhaust Vel15,700 m/s
    High-ε Engine
    Isp1,950 sec
    Exhaust Vel19,100 m/s
    Inert Mass36,000 kg
    Payload Mass
    (incl. crew)
    62,800 kg
    Dry Mass98,800 kg
    Propellant Mass33,620 kg
    Wet Mass132,420 kg
    Mass Ratio1.34
    Exhaust Vel19,100 m/s
    ΔV5,590 m/s
    Initial Accel2.27 m/s
    (0.23 g)

    Remember that εgas is the ratio of power in the 233U Plasma to the total. Hyland's Mini-Gas Core has a εgas of 0.221, or only 1/5th of the power is in the plasma. Beveridge focused on a Low-ε engine with εgas = 0.51 and a High-ε engine with εgas = 0.673.

    Beveridge found that it was not optimal if εsolid is larger than 0.50, that is, if more than 50% of the total engine power comes from the driver core. Hyland's design had εsolid = 0.780, or almost 80%. This means the Hyland's driver core needed more cooling than the cavity wall.

    The obvious solution won't work. Rockets in general use cold propellant to cool off engine components. So one would think the solution is to cool off the driver core with propellant, then send it into the chamber to be superheated by the uranium plasma. But since Hyland's engine only had about 20% of the total energy generated by the uranium plasma, the plasma would not significantly heat the propellant more than the driver core already had. Bottom line is the performance would be about the same as a garden-variety NERVA solid core reactor, but with an engine that was much more expensive.


    In Hyland's engine the driver core produces 78% of the power. Since the driver core is a solid-core reactor, it cannot go above 3,000K or it will melt. Since the uranium plasma is only 22% it can only heat the propellant about 500K more for a total exhaust temperature of 3,500K. Which is about the same as a bog-standard solid-core NTR.

    But if both the driver core and the uranium plasma produce 50% of the power, then the gas core can add about 2,400K more for a total exhaust temperature of 5,400K which is much better than a solid-core NTR. But wait! There's more! Above a temperature of about 5,000K, molecular hydrogen propellant dissociates into monoatomic hydrogen (single-H). This could increase the exhaust velocity and specific impulse by up to a factor of 1.4 (i.e., √2).

    To avoid that unhappy state of affairs, you have to use the un-obvious solution of using a heat radiator to cool the driver core. The trouble is that heat radiators add literally tons of penalty-mass to the engine. You will have to dial down the total engine power to control the heat radiator mass. The end result would be an engine with about the same mass as a standard nuclear-electric propulsion engine (NEP), fractionally more thrust, and drastically less specific impulse (Isp of 2,000 sec instead of 6,000 sec.) In which case it would be more advantageous to use NEP.

    Since neither of those solutions works, Beveridge found a third option. Design the engine so that the driver core power is less than 50% of the total. This means the driver core can be cooled by propellant, and the uranium plasma will most certainly heat the propellant more than the driver core did. A heat radiator is used to cool the chamber from uranium plasma heat. Bottom line: high specific impulse and high power.

    Beveridge did comparison studies on a pure open-cycle gas core, a Low-ε hybrid engine with εgas = 0.51 and a High-ε engine hybrid with εgas = 0.673. For comparison purposes they were all scaled to have a power level of 3 gigawatts. Unsurprisingly the low-ε had the lowest critical mass, the pure open-cycle had the highest, and the high-ε was somewhere in the middle.

    Exhaust Velocity17,658 m/s
    Specific Impulse1,800 s
    Thrust17,800,000 N
    Thrust Power0.2 TW
    Mass Flow1,008 kg/s
    Total Engine Mass127,000 kg
    Uranium Hexafluoride
    ReactorGas Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power1 kg/MW
    Thrust Power0.495GW
    Exhaust velocity22,000 m/s
    Specific Impulse2,200 s
    Thrust45,000 n
    Engine mass66,000 kg
    Diameter5 m
    Length5 m
    Fuel Temp20,000° R
    Propellant Temp10,000° R

    Gaseous core coaxial flow fission / nuclear thermal rocket.

    The basic problem of gas core nuclear rockets is ensuring that the hot propellant escapes from the exhaust nozzle, but the nuclear fuel does not. In this concept, the propellant and fuel are kept separate by a velocity differential. That is, a central, slow moving stream of fission fuel heats an annular, fast moving stream of hydrogen.

    Yes, the uranium jet is aimed straight at the exhaust nozzle. But they figured the uranium loss would be acceptable as long as 25 to 50 times as much hydrogen propellant escapes compared to uranium fuel (measured by mass).

    No, the concept does not work very well. In theory the difference in velocity should keep the uranium/plutonium and the hydrogen separate. Unfortunately the velocity differential at the boundry between the propellant and fuel generates shear forces. The fast hydrogen strips off uranium atoms from the slow fuel plume like a carpenter's plane (laminar and turbulent mixing processes). This means the hydrogen to uranium escape ratio drops below 25.

    The concept seems to have been abandoned.

    This is from Estimates Of Fuel Containment In A Coaxial Flow Gas-Core Nuclear Rocket (1970).

    Again the idea is to have all the furiously hot hydrogen propellant go shooting through the exhaust nozzle, while trying to prevent from escaping as much as possible of the dangerously radioactive and hideously expensive uranium fuel. The point of the paper is to use computer simulations to draw graphs predicting how much fuel will escape given a specific propellant-to-fuel flow ratio. The end result of the calculation is the contained fuel mass (how much fails to escape) in the form of a dimensionless number called the "fuel volume fraction." This is the fraction of the cavity volume occupied by fuel.

    The analysis uses a coaxial free-jet computer model along with custom-made eddy viscosity equations, neither of which they reveal in the paper. They assume a smooth inlet velocity profile. They also assume that the cavity is a cylinder with the diameter equal to the length.

    They plot how the fuel volume fraction varies with different flow ratios, fuel radius, and fluid density. They looked at propellant-to-fuel flow ratios from 10 to 100, fuel-to-propellant density ratios from 1.0 to 4.7, and fuel-to-cavity radius ratios from 0.5 to 0.7. The predicted results more or less agrees with data from previous physical experiments. "More or less" is defined as within ±30%.

    The vaporized uranium fuel stream in the axis is surrounded by the lighter, faster moving hydrogen propellant stream. The coaxial flow should in theory contain the fuel and keep it from escaping even at very high fuel temperatures. In practice though the containment is less than perfect. The large difference in velocity between the fuel and the propellant streams causes turbulent mixing. The goal is to predict the contained uranium fuel mass for various flow ratios (i.e., how much uranium does NOT escape out the tail-pipe).

    You need to know the contained fuel mass:

    • So that the engine can be designed such that the contained full mass is above critical mass. Otherwise there is no nuclear fission and the engine just looks stupid sitting there with sputtering hydrogen flatulence.

    • So that the engine design can be optimized, selecting a engine parameters that push the fuel volume fraction as close to maximum as possible. You don't want to waste uranium, that stuff is more expensive than ink-jet printer ink.

    • So that the engine be designed such that the specific impulse and thrust meets the propulsion system requirements. Otherwise the project boss will be very angry.

    A target engine design would have a desired fuel volume fraction of 0.20 and a propellant-to-fuel flow ratio of 50. The analysis indicates this can be achieved with a fuel-to-propellant density ratio of 1.0 and a fuel-to-cavity radius ratio of 0.7.


    In theory an open-cycle gas-core NTR can achieve a specific impulse greater than 1,500 seconds (exhaust velocity greater than 14,700 m/s) and thrust on the order of 2,000,000 Newtons. The slower-moving uranium fuel stream is at about 55,000°C, the faster moving hydrogen propellant stream is heated by the uranium to about 5,500°C. According to one reference a desirable engine should contain enough fuel to give a fuel volume fraction of at least 0.20 at a propellant-to-fuel flow ratio of 50 or greater.

    A solid rod of uranium is inserted into the engine where is is vaporized by fission heating to form the fuel vapor cloud. The analysis assumes that downstream of plane A-A the fuel is completely vaporized and flowing nearly parallel.

    Flow Model

    Figure 2 shows the mathematical model.

    Continuity equation:

    Momentum equation:

    Mass diffusion equation (Schmidt number assumed to be Sc = 0.7):

    Intitial and boundary conditions are:

    Eddy Viscosity

    For the region near the inlet:

    and downstream

    Having said that, the location of x12 is where ε1 = ε2. With the present calculation the cavity is far shorter than x12, so you can ignore equation (5).

    Inlet Velocity Profiles

    One of the references actually put a plane of porous material at plane A-A to smooth out the inlet velocity. Otherwise the turbulence reduces fuel containment.

    The equation for smooth inlet velocity profile is:

    The report generalizes the inlet velocity profile by using the variable RB (inlet velocity profile half-radius) which they set equal to the upstream buffer radius from one of the references. So RB/RF values (ratio of inlet buffer radius to fuel stream radius) in the reference of 1.14, 1.22, and 1.3 correspond to the radius ratios RF/RC (ratio of fuel stream radius to cavity radius) of 0.7, 0.6, and 0.5.

    Fuel Volume Fraction

    The fuel mass is calculated as the fuel volume fraction VF. This is the fraction of the cavity volume occupied by pure fuel vapor if it were gathered into a central volume at its original temperature and cavity pressure. Cavity volume is defined by planes A-A and B-B, and the streamline through r=RC at inlet (see figure 2).

    The fuel volume fraction is:

    With a known pure fuel density (for a specific cavity pressure and average temperature), VF is a direct meaure of the fuel mass contained inside a full-sized heated engine.


    As mentioned above, the experiental data from the references were for radius ratios RF/RC (ratio of fuel stream radius to cavity radius) of 0.5, 0.6, and 0.7. Also from the references the experimental data for fuel-to-propellant density ratio was from 1.0 to 4.7.

    In the graph, the white circles are experimental data for density ratio 1.0 and the black circles are experimental data for density ratio 4.7. The x-circle is for a desired engine design with a fuel volume fraction of 0.20 and a flow ratio of 50.

    The upper curved line is drawn by the report's equation for density ratio of 1.0. The lower curved line is drawn by the equation for density ratio of 4.7.

    The point of the report is that all but four of the 23 experimental data points fall within ±30% of the equation's curves.


    By using cross-plotting and other clever mathematical tricks, the three figures 7a, 7b, and 7c can be collapsed into one graph: figure 8.

    The report notes that eighty percent of the experimental data points fall within ±30% of the calculated correlating curve. So the equations appear to be close to predicting reality.

    Vortex Confined
    Vortex Confined
    Vortex Confined
    Exhaust Velocity19,620 m/s
    Specific Impulse2,000 s
    Thrust50,400 N
    Thrust Power0.5 GW
    Mass Flow3 kg/s
    Total Engine Mass114,116 kg
    Frozen Flow eff.75%
    Thermal eff.70%
    Total eff.53%
    Uranium Hexafluoride
    ReactorGas Core
    Vortex Confined
    RemassSeeded Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power231 kg/MW

    The hotter the core of a thermodynamic rocket, the better its fuel economy. If it gets hot enough, the solid core vaporizes.

    A vapor core rocket mixes vaporous propellant and fuel together, and then separates the propellant out so it can be expelled for thrust. Energy is efficiently transferred from fuel to propellant by direct molecular collision, radiative heat, and direct reaction fragment deposition.

    The open-cycle arrangement illustrated accomplishes this by spinning the plasma mixture in a vortex maintained by tangential injection of preheated propellant from the reactor walls. The denser material is held to the outside of the cylindrical reactor vessel by centrifugal force. The fuel is subsequently cooled in a heat exchanger and recirculated for re-injection at the forward end of the reactor, while the propellant is exhausted at high velocity.

    The plasma source can be fission, antimatter, or fusion.

    For fission reactions, the outer annulus of the vortex is high-density liquid uranium fuel, and the low-density propellant is bubbled through to the center attaining temperatures of up to 18500 K. A BeO moderator returns many reaction neutrons to the vortex. Prompt feedback actuators maintain a critical fuel mass in spite of the turbulent flow of water or hydrogen propellant. Since the core has attained meltdown, reaction rates must be maintained by fuel density variation rather than with control rods or drums.

    For antimatter reactions, swirling liquid tungsten (about 4 cm thick) is used instead of uranium, for absorbing anti-protons.

    For fusion reactions, it is the propellant that is cooler and higher in density, and thus it is the reacting fuel ball that resides at the center of the vortex.

    N. Diaz of INSPI, 1990.

    From High Frontier by Philip Eklund
    Wheel Flow

    This is from Wheel-flow gaseous-core reactor concept (1965).

    John Evvard figures this gas-core rocket will have (like all the others) an upper limit of about 3,000 seconds of specific impulse, exhaust velocity of about 29,400 m/s. The design is trying to increase the propellant to fuel mass flow ratio to something between 25 and 50. Since uranium has something like 238 times the molecular weight of hydrogen increasing the mass flow ratio is very hard to do.

    The brute-force approach does not work. If you increase the engine pressure to 2,000 psi with a partial-pressure ratio of 80, preventing the reaction chamber from exploding will increase the reactor mass to something between 250,000 to 500,000 pounds. With that penalty weight the propellant load will have to exceed 500,000 to 1,000,000 pounds to capitalize on the increase specific impulse the engine enjoys over a conventional solid-core NTR. And even then the fuel mass flow ratio would be below 25. So this is a dead end.

    So the standard solution is to somehow make an incredibly high hydrogen-uranium volume flow ratio.

    There are numerous schemes to increase the volume flow.

    The vortex-confined GCR makes a vortex of gaseous uranium (sort of a smoke ring) with the center hole aligned with the thrust axis. Hydrogen is injected around the outer edge of the vortex, travels radially across the furiously fissioning uranium being heated all the way, enters the hole in the center of the smoke ring, turns 90 degrees and goes rushing out of the hole and out of the exhaust nozzle.

    The pious hope was that the centrifugal forces acting on the heavier uranium atoms would counteract the diffusion drag of the inwardly moving hydrogen. Sadly the drag produced by the flowing hydrogen is so great that it carries along too much of the valuable uranium.

    The coaxial-flow reactor was another idea that failed even harder. The uranium gas in the center moved really slow while the hydrogen gas around the rim moved really fast. The regrettable result was the velocity difference caused shear forces which allowed the dastardly hydrogen to drag uranium along with it right out the exhaust nozzle.

    John Evvard had a fresh idea: the Wheel-Flow Confined GCR.

    The problem with the vortex confined GCR was that the hydrogen moves through the uranium. This allows the hydrogen to drag along some uranium. The problem with the coaxial-flow is that though the hydrogen doesn't move through the uranium, it is moving at a vastly different velocity. This causes shear forces that allow the hydrogen to drag along some uranium.

    So Evvard tried to find a geometry where the hydrogen does not move through the uranium and it moves at the same velocity as the uranium.

    In the Wheel-flow there is a cylinder of gaseous fissioning uranium in the center of the chamber, spinning around its long axis.

    Hydrogen is injected at the outer surface of the cylinder and moves along the surface, not moving through the uranium. This avoid the vortex-confined GCR's problem. The hydrogen moves at the same velocity as the uranium gas cylinder. This avoids the coaxial-flow GCR's problem. The uranium and the hydrogen rotate as one, as if they were a solid wheel.

    After one rotation of the cylinder the hydrogen is good and hot. It then exits tangentally from the chamber into an array of exhaust nozzles. And there is your thrust.

    Uranium will be lost due to fission and some unavoidable diffusion into the hydrogen. Fresh uranium will be injected from the two end walls, entering the long axis of the uranium cylinder. The end wall will also rotate to match the wheel, to avoid stirring up turbulence.

    The main drawback is that the boundary layer between the hydrogen and uranium is unstable. Any blob of uranium entering the hydrogen blanket will be accelerated outward by simple boyancy. This could possibly be stablized by an axial magnetic field. The fissioning uranium is more ionized than the hydrogen so the magentic field will grab the uranium more firmly.

    Since the temperature inside the reaction chamber is hot enough to vaporize any material object the ions are moving like microscopic bats from hell. You'd think the high uranium molecular velocities would make the uranium cloud instantly explode to fill the chamber. Luckily the mean free path of individual atoms is a microscopic 10-7 meters or less (one micrometer, about the length of a bacteria). Since the hot uranium atoms cannot move further than the span of a typical e coli germ without crashing into other atoms their effective speed is slowed down about the same as the wheel rotational velocity.

    The report is a little vague about this design. It says that if the wheel-flow engine is used in a gravitational field, the spinning cylinder of fissioning uranium might settle to the bottom of the chamber, which is bad. However, unless you were using an open-cycle gas core nuclear engine spraying radioactive death from the nozzle as an aircraft engine I don't see the application.

    The idea seems to be that while some hydrogen is injected around the uranium gas cylinder for coolant, most of the propellant hydrogen goes across the top along the axial line. I guess the propellant lowers the gas pressure enough to levitate the uranium cylinder.

    In the standard wheel design, the end walls will have to be cooled since they are exposed to the fury of fissioning uranium. This can be avoided by bending the uranium gas cylinder so the ends meet, converting the cylinder into a torus donut shape. Since it is now a ring there are no end walls and no need to cool them.

    The problem is that the end walls were where the fresh uranium was injected, and it is unclear how to refresh the torus.

    This design makes a bit more sense. It uses a torus of uranium gas. The rocket rotates around the thrust axis to make artificial gravity. This pulls the torus outward, making it expand. Meanwhile the propellant hydrogen is roaring down the thrust axis, being heated and expelled out the exhaust nozzle. This lowers axial gas pressure and pulls the torus inward, making it contract.

    Between the artificial gravity and Bernoulli's principle the torus of uranium is held in place.

    Of course there is still the unsolved problem of how to refresh the torus.

    MHD Driven Rotation
    MHD Driven Rotation
    T/W ratio1:1211:1121:135
    mass flow ratio
    Mass flow
    Radiator area
    Fission Power
    Mass Schedule
    (metric tons)
    Cavity Struct1246942
    Cryogenic Magnet12898128
    Turbo-electric Gen999634482
    TOTAL ENGINE5,6904,0104,110

    This is from Gas-Core Nuclear Rocket With Fuel Separation by MHD-Driven Rotation (from Research on Uranium Plasmas and their Technological Applications page 155) (1970)

    The major problem with open-cycle gas core NTR is keeping the blasted uranium from escaping out the exhaust nozzle. Or at least only escaping after all the expensive uranium has been burnt in nuclear fission.

    The researchers noted that injecting gas tangentially at high velocity (Figure 1 (a)) would confine the U235 fuel to the outer region of the cavity yet allowing the propellant to diffuse radially into the center region was the great hope. The supersonic rotation would develop high centrifugal force, forcing the heavy U235 to the outer while permitting the light hydrogen into the middle. Nope, this was tested and it don't work no-way, no-how (though they put it "this configuration does not result in effective separation of the two gases").

    The researchers noted the next great hope was vortex-stabilization. The idea here was to exploit the high stability of the rotation flow instead of centrifugal force. The flow was subsonic, so the centrifugal force was negligible. A donut-shaped vortex of U235 floats in the center of the chamber, while hydrogen propellant flows around the edge of the chamber. The idea is that the hydrogen will stay in the outer regions due to the inherent stability of the rotation flow.

    This didn't work either.

    Looking more closely at the scheme of injecting gas tangentally, it was noticed that the centrifugal separation worked best in the part of the flow that resembled solid-body rotation. That is, rotating as if it was a solid brick of matter instead of rotating gas. So investigation focused on making the U235 and hydrogen gas rotate as if it was a solid body.

    The trouble with tangential gas injection is it develops into what is called an "inviscid vortex flow field". I don't know that that is either, but apparently it makes TGI about as effective as trying to push spaghetti uphill, when it comes to making a centrifuge. So a different means of rotating the gas will be needed.

    In Figure 1 (b) the entire cavity chamber will be spun mechanically to induce the centrifugal effect. Alas, in order to separate the uranium from the hydrogen you need about 1,000,000 gs of centrifugal force. No known construction material can withstand that sort of force so the chamber would explode like a bomb. To keep it from exploding you'd have to apply an external gas pressure of at least 10,000 atmospheres, which is rather excessive. What's worse is the friction loss from the external gas would become prohibitively large. Another failed concept.

    The study authors had an idea. How about spinning the uranium and hydrogen using electromagnetic forces? See Figure 1 (c)

    This uses the magic of Magnetohydrodynamics.

    If you send electrical current J in a direction parallel to the thrust axis (z-axis), a magnetic field B will be created radially at 90° (r direction). This is called J×B. The important point is the magnetic field will be pushing the gas towards the chamber walls, allowing the gases to rotate as if it was a solid body. Rotating such that the blasted U235 separates from the propellant, and is kept away from the exhaust nozzle.

    No chambers rotating so fast they explode, no friction loss from external gas at 10,000 atm, this design looks like it would actually work.

    The report presents a prototype engine.

    The cavity is typically seven to eight meters in both diameter and length. The radial mangetic field is produced by cryogenically cooled magnetic coils, and has a strength of about 1.0 tesla near the cavity walls. The coils draw a negligible amount of power. The total electrical current is about 20,000 amps, flowing through 200 pairs of segmented electrodes. The chamber pressure is 20 to 60 atmospheres. The maximum tangential rotational velocity is about 1.6 to 1.8 kiometers per second.

    The hydrogen propellant is introduced along the centerline of the cavity and flows through the center at low velocity. The propellant is heated mainly by radiation, plus a bit by conduction on the part of any propelant that brushes the hot uranium. Hydrogen propellant is regrettably mostly transparent to infrared heat rays, which is why most gas core designs seed the hydrogen with microscopic tungsten particles or something. In this design they seem to be relying upon uranium atoms for seeding, since some will diffuse into the propellant (about 0.00015 mole fraction).

    The chamber temperature varies from 10,000K at the chamber walls (in the hot uranium) to 6,000K at the centerline (in the propellant). This averages out to a specific impulse of approximately 1,770 seconds, which is actually pretty good. The thrust is about 40 metric tons or 392 kiloNewtons.

    The liquid hydrogen first flows through the superconducting magnetic coils to help keep them at cryogenic temperatures. It then passes through a pump to pressurize it to about 640 atmospheres. Now the propellant travels through the moderator-reflector wall heat exchanger, simultaneously cooling it off and getting real hot. The hot hydrogen enters the first-stage turboelectric generator, which supplies half the required MHD power. The hydrogen leaves the turbogenerator at about 107 atmospheres of pressure. It then passes through the moderator-reflector heat exchanger a second time, and enters the second-stage turboelectric generator. This produces the other half of the required MHD power. Finally the propellant enters the chamber to be superheated and exhausted to create thrust.

    The propellant only removes about 3% of the waste heat in the moderator-reflector. The other 97% of the waste heat is removed by a liquid-metal cooling system which expells the heat out of a heat radiator with a surface temperature of 1,200K. A fraction of this is diverted to an auxiliary power generator for other power needs.

    RD-600: Soviet bimodal GCNTR
    RD-600 bimodal GCNTR
    Thrust5,880,000 N
    Specific Impulse2,000 sec
    Chamber Pressure500 kg/cm2

    This is astonishing. A rocket expert living in Russia, Denis Danilov, stumbled over this in their research. This is a proposal for a Soviet gas core nuclear thermal rocket project. That ran from 1963 to 1973. And was bimodal!

    The fact that the Soviet Union actually had a real live gas core project that ran for ten years and employed 90 researchers got my attention.

    The fact it was a gas-core BIMODAL engine made my jaw drop. I have never ever seen a proposal for a bimodal nuclear engine that used anything other than a solid-core nuclear engine. According to this document (page is in Russian, use Google Translate if need be), in power generation mode it uses a separate circuit for the uranium which has no connection with the outside world. Yes, this adds more points of failure, but on the plus side it allows one to generate power without spraying fissioning uranium out the exhaust like it does in thrust mode.

    While in thrust mode there is a second MHD power generator around the exhaust nozzle, to harvest a bit of thrust to make electricity. In coast mode the uranium is diverted from the rocket engine altogether, entering a closed cycle which energizes the main MHD generator.

    I did spot a single sentence in one of the documents that seems to imply an important advantage to gas core MHD power generation.

    In the following documents:

    • Type A NTR: NERVA style solid core nuclear thermal rocket
    • Type V NTR: open-cycle gas core nuclear thermal rocket
    • tf: Tonnes-Force. 1 tf = 9806.65 Newtons

    This is from a thread in the Kerbal Space Program forum by DDE:

    Dear Mr. Chung,

    I have been a user of your website for quite a while. However, I’ve stumbled onto a virtually unknown piece of Red Atomic Rockets history that I’d like to share with you. I’ll stick mostly to direct translation of sources to avoid putting my spin onto it.

    So, there I was trawling through the full list of Energomash rocket engines at, trying to make sense of their classification scheme. I knew RD-1xx were kerolox, RD-2xx were hypergolic, RD-3xx involved fluorine and had to be given a wide berth, RD-4xx were early, solid-core NTRs, RD-5xx used peroxide, and RD-7xx switched from kerolox to hydrolox on the fly to maximize total Δv. I was trying to find out what the heck an RD-6xx was, thinking it may have been the bimodal NTR branch.

    There was only the RD-600. What has a vacuum thrust of 600 tonne-forces (5,880,000 N), an Isp of 2000 sec and chamber pressure of 500 kg/cm2?

    I was hooked already, and the Russian Wikipedia, messy as it is, answered me with an article about the solid-core twisted-ribbon RD-0140 ( that contained a timeline apparently copied from, which is a webpage of the Department of physical mechanics of the Moscow physico-technical university. Translation follows:


    To produce a gas-core nuclear rocket motor with high specific impulse (>3000 sec) for missions to planets of the Solar system.


         1957 – Initiation of research as suggested by V.M. Ievlev and approved by I.V. Kurchatov, M.V.Keldysh and S.P. Korolev (by then known as the Three Ks as a result of their ICBM work – ed.)
         1953 – Government decree on research into “cruise missiles propelled by ramjets exploiting nuclear energy”
         1955 – formation of research group at Air Industry Ministry NII-1 (currently Keldysh Research Centre – ed.). Led by V.M.Ievlev (K.I.Artomonoc, A.S.Koroteev et al). Objective: development of NTRs “type A” (Isp=850-900 sec) and “type V” (up to 2000 sec).
         1956 – Government decree on “creation of a long-range ballistic missile with an atomic engine”. Chief designers: overall – S.P.Korolev, engine – V.P.Glushko, reactor – A.I.Leypunskyi; personnel recruitment and training at the Moscow Aviation Institute – N.N.Ponomarev-Stepnoi.
         1958 – Government decree on initiating NTR research; overall command relegated to M.V.Keldysh, I.V.Kurchatov and S.P.Korolev
         1958 – construction of reactor test stand and “hot lab” commences at Ministry of Defence Proving Grounds №2 (Semipalatinsk nuclear testing site)
         1964 – combined Central Committee of the Communist Party of the Soviet Union and Government of USSR decree on initiating construction of the Baikal firing complex at the Semipalatinsk NTR test site.
         1966 – creation of 11B91, an A-type NTR (non-GRAU designation RD-0140 – ed.). Scientific supervision – Keldysh Centre (V.M.Ievlev), manufacturing – Chemical Automatics Design Bureau (A.B.Konopatov), fuel elements – Perm research and technological institute (I.I.Fedik)
         1968 – development of RD-600 GCNR; scientific supervision by the Keldysh Centre, lead design by NPO Energomash under V.P.Glushkon; thrust 6 MN, Isp 2000 sec
         1968 – Government decree on development of RD-600 GCNR and construction of the Baikal-2 test stand
         1970 – NPO Energomash and the Keldysh Centre complete a draft proposal for a 3.3 GWt gas-core powerplant, EU-610
         1972 – criticality of an IVG high-temperature research reactor at Baikal (N.N.Ponomarev-Stepnoy)
         1978 – criticality of the first 11B91 NTR, Image caption: high-temperature GCNR variant (“Type V”)


         1955 – commencement of work on a Type A NTR (SCNR) at Los-Alamos under the Rover program
         1960 – conceptual development of a Type V NTR (GCNR) by Weinstein, Kerrebrock (MIT) and Los-Alamos; Isp=(600-2000) sec
         1963 – development of Nuclear Engine for Rocket Vehicle Applications (NERVA) by Westinghouse and Los-Alamos
         1962-1968 – experiments in hydrodynamics, plasma stability, heat physics and radiation of uranium plasma, optical qualities of hydrogen, neutron calculations of reactor criticality
         1973 – cessation of TR research



         1985 – Los-Alamos and NASA conduct systemic analysis of Lunar missions, concluding that resumption of research on Type V systems is crucial (twofold cost and flight duration reduction). Equipment and systems are preserved in Los-Alamos and Nevada (Keldysh Centre and Semplatinsk).
         1989 – President Bush announces the Space Exploration Initiative – a manned mission to Mars by 2018 (see Russian space program). NTRs assumed as baseline by NASA and Los-Alamos. DOE/NASA task force on NTRs formed.
         1991 – GCNR conference in Los-Alamos
         1992 – research into stability, neutrons, displacement, quantitative modelling, MHD (presumably, magnetohydrodynamic generators or magnetohydrodynamics in general – ed.)
         2005 – China and Kazakhstan declare research into spaceborne nuclear power a priority

    The GCNR program in the US has been unsuccessful due to “lack of experimental data on thermophysical qualities of substances and the calculating power for modelling high-temperature hydrodynamics and turbulence (from MIT report by R.Patrick & Kerrebrock). The USSR has resolved these issues with participation of the Department of physical mechanics

    Both the US and USSR would relegate large rocket projects to a triad of R&D Centre-University-Test Ground, e.g. Los-Alamos-MIT-Nevada and Keldysh Centre-Moscow Institute of Physics and Technology-Semipalatinsk

    Key development aspects:

    • Handling and operation of a gaseous fuel element
    • Thermophysics of nuclear fuel and reaction mass
    • Vortex and magnetic hydrodynamics
    • Radiative and convective heat and mass exchange
    • Thermal protection of reaction chamber walls and the egress canal
    • Achieving GCNR criticality
    • Achieving stable GCNR operation despite high mobility of fission fuel

    GCNR parameters:

    • Pressure – 1000 atm
    • Temperature: fuel 40-60 thousand K, reaction mass 8-10 thousand K
    • Molten uranium at 1500-2300 K
    • High-pressure hydrogen at up to 2800 K
    • Chemically aggressive environment due to alkali metals at up to 2800 K
    A.S.Koroteev, E.E.Son. Development Nuclear Gas Core Reactor in Russia: AIAA-2007-0035 (behind paywall), 45th AIAA Aerospace Sciences Meeting and Exhibit, 2007, Reno, Nevada.

    By then I was thoroughly intrigued and began to go up and down Yandex search results while trying to filter out the mentions of the RD-600 turboshaft engine. I’ve stumbled upon the questionably legal (as is most of Russia’s internet) online version of a collection of Glushko’s works, published by Energomash in 2008 with a total run of 250 books, that contains a few interesting official documents of his own authorship.

    August 18 1963

    to the research and development council of the USSR State commission on defence technology
    on prospective R&D at OKB-456

    The research and development conducted at OKB-456 have led to the following conclusions:

    1. Further development of oxygen and RFNA engines at OKB-456 is inexpedient, as most of the expected pathways of rocket development are better served by other engine types. This does not mean that oxygen and RFNA motors are undeserving of further development, as occasionally they are highly fit for purpose. For example, when performance takes a back seat to propellant liquidity range, RFNA engines are quite applicable. Liquid oxygen, non-toxic and cheap, is also quite successful, despite difficulties associated with its low boiling temperature…

    (sections 2, 5, 10, 11, 15 missing – ed.)

    3. At the present stage the development of high-powered liquid propellant rocket motors for surface-to-surface missiles and space boosters at OKB-456 relies entirely on UDMH and N2O4.

    The use of storeable, hypergolic fuel components already mastered by the chemical industry has ensured the development of highly effective motors for R-36, 67S4, R-56 and the first stage of UR-500 (a.k.a. original variant of Proton – ed.) in accordance with the Decrees of CC of CPSU and Government of USSR.

    These engines range, in thrust from 12 (vacuum) to 600 (sea level) tf, and in specific impulse at sea level from 272 to 300 sec, in vacuum up to 325 sec.

    4. OKB-456’s current plan outlines the development of a gas-core nuclear motor with liquid hydrogen as reaction mass, with thrust of 200-600 tf (RD-600) and specific impulse of 2000 seconds.

    The creation of such an engine would be a true revolution in rocket science due to the dramatic leap in specific impulse.

    Further development of this engine class can allow, with time, specific impulses in the 2500 sec range, allowing creation of booster rockets with an order of magnitude greater payload than regular chemical-powered ones.

    6. For boosters intended for first or second space velocities (orbital and planetary escape trajectories respectively – ed.) the optimal design is as follows: the chemical motors on the first stage loft the second stage carrying the GCNR to the minimum safe altitude dictated by the contamination of the exhaust by fission products…

    7. The low-altitude loft stage in the abovementioned GCNR-based system should rely on the use of chemical rockets using high-density storeable propellants (UDMH and N2O4), as under such conditions the fuel is more energetically efficient than kerolox and, due to its chemical stability and hypergolicity, more convenient. An UDMH-N2O4 first stage can achieve Isp of 300-320 sec and a mass ratio of 1.18 (e.g. 8D420 engines).

    8. Use of a GCNR makes the development of an entirely reusable booster more realistic by allowing the use of an airbreathing first stage…

    9. In case of two-way missions to the planets and their moons using a GCNR the payload mass can be further increased by using a third stage with refireable electric rockets with Isp in the 10000-20000 sec range. (here and henceforth bold added as emphasis – ed.)

    12. (sales pitch for high-energy, storeable lander propellants, including UDMH-N2O4, late RD-5xx series motors burning H2O2 and pentaborane, and the RD-550 using H2O2 and beryllium hydride (!); sly suggestion to cease development of all cryogenic chemical rockets – ed.)

    13. The preceding deliberations on GCNR application can only be realized once there is confidence of the possibility of creation of said motor, backed by experimental research, including stand tests of a single fuel element motor with a gas core reactor, in a state approaching operational (plasma temperatures up to 30000 K, pressure up to 500 atm)…

    14. The prospective development plan of OKB-456 has been developing for the last few years along the lines laid out in sections 1-13 and currently ranges out to 1970. In accordance with it:

    d) Key developments are

    I) High-powered UDMH-N2O4 motors: for first stages of R-36 and 67S4 (8D723, 8D724), for first stages of UR-500 and R-56 (PD43), for second and third stages of R-56 (11D44, 8D724) and a high-performance motor with ASL thrust of 600 tf and specific impulse of 300 sec ASL and 323 sec in vacuum for the first stage of heavy boosters (8D420);

    II) Upper stage motors of 10-12 tf using: UDMH-N2O4 (8D725, Isp=325 sec), H2O2-pentaborane (11D11, Isp=375 sec), H2O2-beryllium hydride (RD-550, Isp=400-460 sec, under investigation);

    III) Gas-core nuclear rocket with liquid hydrogen reaction mass with thrust 200-600 tf (RD-600, Isp=2000 sec) as second-stage engine…

    16. (section on weaker chamber pressure in US chemical rockets leading to poorer performance at comparable mass – ed.)

    OKK-456 chief designer, academician GLUSHKO

    Archive 1727 (123-130)

    Backtracking a bit, I’ve found an earlier memo containing some of the points missing from the above report

    May 6 1963


    9. The above deliberations on expedient applications of liquid-fuel rocketry using storeable and cryogenic propellants, solid-core and gas-core NTRs and electric rockets can only be realized once the creation of GCNR is proven feasible in the coming years, backed by experimental research, including stand tests of a single fuel element motor with a gas core reactor, in a state approaching operational (plasma temperatures up to 30000 K, pressure up to 500 atm). Creation of the testbed facility and an experimental single fuel element motor is planned for late 1965, with tests commencing in 1966. First experimental results are likely to be produced by 1967. Should development outcomes prove favourable, a flight-ready GCNR (RD-600) can be deployed by 1970.

    OKK-456 chief designer, academician GLUSHKO

    Archive 1727 (66-71)

    Here’s the economic aspect:

    April 29 1969

    comrade ABRAMOV I.I.

    Re: development of Type V NTR

    In accordance with the Decree of CC of CPSU and Government of USSR 524-215 of June 19 1964 the Ministry has initiated the program “Development of RD-600 nuclear rocket motor” at an estimated cost of 20 million roubles.

    Over 1960-1968 Energomash and related organizations have conducted a range of design, theoretical, production and experimental studies the results of which are found in the following reports:

    1. Draft project of RD-600 GNCR (original sent to MGM on September 30 1964)
    2. Report on preliminary project of a testbed loop-type motor with a single gas-core fuel element (original sent to MGM on April 9 1965)
    3. Report on progress on RD-600 NTR for 1966 (original sent to MGM on December 31 1966)
    4. Report on progress on RD-600 NTR for 1967 (original sent to MGM on January 4 1968)
    5. RD-600 nuclear rocket motor. Abbreviated results of theoretical and research studies (original sent to MGM on December 18 1968); said report is equivalent in content to a complete pre-draft project; it has been approved in that capacity by Central Research Institute of Machine Building in correspondence of April 15 1969.

    Over the process of developing the RD-600 a phased approach to development of Type V nuclear rocket motors has been determined to be expedient, including the initial development of a testbed motor with a single gas-core fuel element, alongside the requisite testing facilities.

    The above-mentioned phased approach is approved by the Decree of CC of CPSU and Government of USSR №388-146 of May 24 1968. In particular, the decree outlines the development of a high-performance on-board powerplant based around the single gas-core fuel element design. The technical objective set by the Central construction bureau of experimental machinebuilding (currently RKK Energiya, earlier KB-1, the late Sergei Korolev’s outfit – ed.) calls for a powerplant developing 3 mln kWt (sic), which is achievable by using a reactor of a Type V NTR with a thrust of approximately 50 tf.

    Due to the above I request your permission to close the program titled Development of RD-600 nuclear rocket motor”, to disburse the actual expenses of 9325 thousand roubles, and to commence a program titled “Development of an experimental testbed reactor with a single gas-core fuel element and a draft project of a high-performance powerplant” in accordance with the draft project file attached to my letter of April 1 1969, with an estimated cost of 2416 million (sic) roubles.

    Chief designer GLUSHKO

    Archive 82/125 (36-37)

    One of the last documents in the collection is an overview of Glushko’s entire NTR business at the height of its glory. Believe me, most of the more informative materials on Soviet rocketry, such as Gubanov’s memoirs about Energiya-Buran, are this dry and technical.

    July 26 1973

    Outline of development of nuclear rocket motors at KB Energomash

    Increase of the specific impulse, being one of the cardinal directions of rocket motor design, has driven the effort to exploit the energy of nuclear fission in this rapidly developing field of technologies. The successes of national rocket design in 1950-1955 have allowed the Physico-energetic institute of the Ministry of medium machinebuilding (cover name for Soviet nuclear armament and energy agency; similar to the Ministry of general machinebuilding seen earlier – ed.) to put forward the concept of integrating a solid-core nuclear reactor into a rocket engine. Based on PEI’s proposal, relying on a uranium-graphite reactor heating up hydrogen, KB Energomash began ongoing research work on solid-core nuclear rockets (Type A) in 1956. Research in cooperation with PEI over 1956-1958 revolved around neutron flux and thermal calculations and covered a variety of thrust levels (tens to hundreds of tf), reaction masses and reactor types (by moderator material, fuel distribution, et cetera). In 1958 Energomash formed a permanent design taskforce focused on nuclear motors, evolving in 1961 into a full-fledged NTR design section. From the start the unit was led by R.S.Glinik; some of the first members ncluded E.M.Matveev, G.L.Lioznov, V.Ya.Sirotkin, K.K.Nekrasov and V.N.Petrov. Based on research by PEI and Energomash along with parallel investigation of rocket propulsion through fission energy by the Keldysh Centre, in 1958 the CC CPSU and Government of the USSR issued a joint decree that, among others, outlined he development of a draft design of a high-thrust Type A NTR using ammonia as reaction mass. The draft design was completed in 1959 with assistance from PEI (neutron flux) and the Keldysh Centre (thermal physics, fuel rod design, engine dynamics research), yielding two variants, RD-401 with a water moderator and RD-402 with a beryllium moderator. RD-402 produced superior performance, at 168 tonne-force vacuum thrust, 428 sec Isp, weight to power ratio 22 kg/tf of thrust (note that the original refers to it as specific power – ed.) and a length of 6760 mm.

    The draft proposal presented the following innovations: proof of applicability of heterogenous multi-fuel rod reactors for thrust up to 500 tf; application of a solid beryllium moderator; fuel rod design providing high reaction mass temperature (up to 3000 K) with minimal fluctuations along the cross-section of the fuel element; homogenized reactor control system using absorber gas in isolated canals; closed-cycle turbopump feed, with the working fluid heated by dedicated fuel rods; steering via gimballing the engine; multi-nozzle design dramatically shrinking the length of the motor; mounting of the booster pump in the outboard section of the rocket’s tankage; proof of the necessity of cooling most of the components due to atomic radiation. Although overall the energetics of the RD-402 were relatively unimpressive, the design and development allowed many of the complex issues in construction of Type A NTRs to be exposed and solved.

    By 1960 positive outcomes of use of liquid hydrogen in rocket design and the availability of requisite tankage technology have allowed to push forward with the draft design of a Type A NTR using hydrogen reaction mass; the CC CPSU and Government of the USSR authorized it in 1960. As a result by 1962 Energomash in cooperation with a range of other units, under general leadership of the Keldysh Centre and with reactor design input by PEI, developed the draft design of the RD-404.

    RD-404 was rated for 200 tf vacuum thrust, 950 sec specific impulse and weight-to-power ratio of 45 kg/tf of thrust including radiation shielding for the tankage, and was 7770 mm long.

    RD-404’s design included a number of design solutions developed specifically for Type A NTRs. After researching a range of moderator options, the final design used an optimized arrangement of beryllium into autonomous cells mated with the fuel rods. A modified sectioned design of the fuel rod was substantiated and implemented, with the initial graphite-coated zone and a post-heat section coated in a metal-carbide composite leading to average hydrogen temperatures in the 3000 K range. A liquid control “rod” system using mercury was conceived and researched. An engine control system relying on specialized pyroautomatics was also developed.

    A complex study was conducted in order to optimize weight-to-power ratio while also developing rocket-engine system design. Research into start cycle and throttling was performed, along with the terminal stage and shutdown cycle using the remnant heat of a subcritical reactor; reverse systems were developed to minimize post-shutdown thrust.

    Engine subsystem design, tankage protection and engine-to-rocket coupling were investigated with regards to reactor radiation effects.

    A steering system using vanes on nozzle edges was also tested and implemented. Initiation calculations very verified by a set of experiments on fuel rod material resistivity, component durability, physical assembly tests, gas-dynamic tests of the multi-nozzle design, tests of the mercury control system, et cetera.

    A considerable contribution to Type A NTR design was on part of a Keldysh Centre taskforce led by V.M.Ievlev and including K.I.Artamonov, R.B.Akopov, V.N.Bogin, A.I.Gori, V.A.Zaitsev, G.V.Konyukhov, E.P.Terekhov.

    The development of RD-401, -402 and -404 essentially led to establishing the key principles of Type A NTR and subsystem design, along with many aspects of manufacturing, testing and operation. Over the course of these projects the organizations involved brought up new creative collectives while laying the groundwork for extensive industrial cooperation. (standard canned Communist phrases, if you can’t tell. – ed.)

    The resultant draft projects of RD-401, -402 and -404 were investigated by a number of highly competent expert panels and technical councils, and were rated highly by them.

    Based on existing experience, in 1962-1963 a draft design was carried out for a mid-range NTR referred to as RD-405, with 30-40 tf thrust, liquid hydrogen reaction mass, specific impulse of 900-950 sec, weight-to-power ratio 55 kg/tonne-force of thrust including tankage shielding. The reactor used a zirconium hydride moderator, beryllium reflectors and fuel rods similar to those of RD-404.

    In actual operation requiring multiple firings the average specific impulse of a SCNR will be degraded by several tens of seconds due to lengthy reheating from standby state.

    Limited possibility of any further increase of specific impulse in Type A NTRs has been the key reason for cessation of research at Energomash in 1963 and relegation of this development topic to a different OKB (see: RD-0140 – ed.)

    It was decided that Energomash would instead focus itself on the much more promising gas-core NTR (Type V), which could lead to a revolutionary leap in aerospace design.

    The research at the Keldysh Centre in 1958-1963 under the overall supervision of V.M.Ievlev, covering principal schemes of gas-core reactors, gas-core fuel elements and Type V NTRs overall, had substantiated the plausibility of a highly energetic engine design, which justified initiating of design development at the construction bureau. A significantly greater specific impulse of Type V designs compared to Type A, making possible the design of spacecraft with qualitatively new capabilities, along with a number of operational advantages (absence of fissionables in the engine during manufacture, possibility of fissionable removal after tests, thus making in-situ servicing and refueling more plausible) make Type V systems exceptionally promising, despite the obvious challenges and the considerable development costs.

    From the very beginning of work on Type V NTR at Energomash in accordance with the Decree of CC CPSU and the Government of the USSR under general scientific supervision of the Keldysh Centre had two key directions: development of an actual high-thrust rocket motor, and the development of a testbed motor used to test off the key functioning principles of the full-scale engine, with a range of theoretical and practical problems being combated along the way and a dedicated test facility being constructed. It should be noted that the development of Type V engines benefitted immensely from the expertise accumulated when developing Type A NTRs.

    The RD-600 was designed in 1964-1968, rated for 600 tf of thrust at 2000 sec specific impulse, with weight-to-power ratio of ~100 kg/tonne of thrust including shielding, and a total length of 14000 mm. The reaction mass is liquid hydrogen doped by lithium. The engine includes a multi fuel element gas-core reactor with a solid moderator and reflector (beryllium, beryllium oxide, graphite), gaseous fuel elements with a central moving stream of fission fuel, and a closed circuit for nuclear fuel, complete with a condenser, separators, pump and fission products removal system. Stabilization of flow in gaseous fuel elements is performed by magnetic solenoids powered by a unipolar electric generator. Research conducted by the Keldysh Centre, PEI and a number of other institutes covered a wide array of topics around designing the principle layout of the motor and optimization of key aggregates and systems in an effort to maximize the specific impulse. Research included neutron flux measurements on physical assemblies, modelling of parallel flows, studying the effects of longitudinal magnetic fields on flow of conductive media, investigating radiative heat transfer and thermal protection of the structure from high-intensity heat, studying resilience of various construction materials when immersed in liquid uranium, seeking production and testing methods of temperature-resistant porous materials, et cetera.

    The executed research proved the principal possibility of producing an NTR with uniquid qualities through use of a gas core reactor, while revealing the extraordinary challenges of organizing the operation, requiring new and specialized materials and manufacturing technologies, associated with such an engine. It was proved that a range of experiments simulating the key operational processes would be crucial for the development of a GCNR, but would require specialized testbeds and a reactor stand facility.

    Creation of such a facility and a testbed motor was outlined by a Decree of CC CPSU and the Government of the USSR of 1968, and had taken centre stage in the work of Energomash and its associates since 1964; it is even more important now, as the theoretical solutions in gas-core reactor design take a backseat to the need for practical testing.

    In accordance with the aforementioned decree in 1970 another design study was completed, this time of a spaceborne electric powerplant dubbed EU-610, with an electric output of circa 3.3x106 kWt, specific power 0.7x105 kWt/kg/sec (sic), relative mass 18.7 g/kWt, length 10000 mm. The powerplant is based around Keldysh Centre’s proposed improvements to the gaseous stream fuel element design. A significant increase in magnetic field intensity coupled with special endcaps have enabled the creation of a “dead space” for fuel cooldown, allowing he creation of a reactor with a single fuel element, with an order of magnitude less power than the RD-600, and without a circulation contour for fission fuel. Further draft proposals outlined the use of this reactor core as a standardized design, producing an NTR with a thrust of 50-60 tf.

    High parameters of hydrogen plasma heated up by a GCNR make it especially appealing for powerplants that use high-efficiency heat-to-electricity conversion using magnetohydrodynamic generators. A draft proposal for such a system was drawn up by Energomash, the Keldysh Centre, PEI and IPPE.

    Over the development of the draft a significant volume of testing of key processes on model systems and of neutron flux profiles on physical assemblies were performed.

    A major role in theoretical and calculative work in designing the Type V system and buttressing the experimental developments was played by I.M.Ievlev and the laboratory under his leadership, namely K.I.Artamonov, N.N.Borisov, A.Ya.Goldyn, A.I.Gorin, M.M.Gurfink, A.M.Kostylev, V.N.Krylov, H.H.Kuznetsov, V.M.Matyshin, A.V.Moskolyov, O.I.Novoznov, A.B.Prishleptsov, A.A.Pavelyev, S.S.Preobrazhensky, E.P.Terekhov, R.A.Fedotov, A.A.Shirokov.

    The development of the EU-610 powerplant will unlock considerable capabilities in achieving astronautic and general fusion power objectives.

    Nuclear rocket motors and nuclear powerplants of Type V, possessing quantitatively and qualitatively greater capabilities compared to chemical and Type A nuclear rockets, are intended to ensure further progress in rocket and aerospace technology development.

    Glushko V.P., Glinik R.A.

    Archive 82/179 (72-80)

    And finally, here’s an essay from a defunct… site, credited to an Aleksandr Valeryevich Khoroshikh at horoshih-aleksander at; it has images, citations, technical details, and it offers a saddening finale to the whole story.

    Gas-core nuclear rockets

    The idea of using a nuclear motor in a rocket dates back to the dawn of that field [1]. An ammonia NTR produced a specific impulse comparable or superior to hydrogen-oxygen chemical motor [2], without requiring sophisticated cryogenics and bloated tankage to contain the low-density hydrogen. In particular, there was a program to develop an R-7-style missile, with an NTR sustainer stage surrounded by six kerolox strap-ons [2]. (see – ed.)

    At the time it was assumed that a nuclear motor would become the core of a successful intercontinental ballistic missiles, ensuring considerable funding of that sphere. However, once chemical motors (especially ones using heptyl-amyl (official codenames for UDMH-N2O4 – ed.)) achieved performance that satisfied the military, the concept of an NTR-based ICBM was abandoned. However (sic), the USSR initiated a space program that included a manned Lunar, and later a Martian mission. One should not forget Nikita Khrushchev’s announcement of the Soviet moonshot. This reinvigorated NTR development.

    An ICBM could only utilize a solid-core NTR, with all other options (liquid or gaseous fuel in a hollow reaction chamber) inevitably leading to some of the fuel escaping and contaminating the environment [4] Furthermore, as mentioned above, reaching America turned out to be within the capabilities of chemical rockets. This relegated NTRs to the single role of a high-efficiency motor for upper stages of boosters and for interplanetary craft.

    Interplanetary missions are especially dependent on engine specific impulse, as the requisite Δv approaches tens of kilometers per second. In this aspect, gas-core NTRs are particularly outstanding, capable of exhaust velocities comparable to electric rockets [5] while also developing thrust comparable to chemical motors. Unlike electric rockets, they can achieve the requisite speed in comparatively short order, rather than months. This in turn allows rapid transit through the Earth’s radiation belts, subjecting the crew to much less radiation. Also, quite notably, not only the minimum energy Hohmann transfers, but “fast-track” trajectories, like parabolic orbits, become possible.

    “The decision to develop NTRs and spaceborne nuclear electric powerplants based on gas-core nuclear reactors was formulated by Energomash chief academician V.P.Glushko in 1963 and was later approved by a Decree of CC CPSU and the Government of the USSR. By then the scientific corps at Energomash had six years of experience in designing and developing SCNRs. Theoretical research into GCNRs had been conducted since 1957 under the leadership of a USSR Academy of sciences corresponding member V.M.Ievlev at the Scientific institutes of heat processes (later the Keldysh Centre). Only two countries, USSR and USA, have attempted to tackle this technology, comparable in complexity to controlled thermonuclear fusion and requiring colossal financial expenditures,” as [6] describes the beginning of GCNR research.

    Energomash’s lead unit on gas-core reactors and derived NTRs was a section led by R.A.Glinik. Solving the problem involved in design required cooperation with numerous institutes (primarily from the aerospace and nuclear industries) and the country’s leading universities under the overall scientific leadership of the Keldysh Centre. Considerable support was lent from the country’s leading scientists, such as academicians M.D.Millionschikov, A.A.Bochvar, Ye.P.Velikhov [6]. One of the key participants was B.I.Katorgin, who was also involved in development of RD-560 (H2O2 and beryllium hydride) and RD-600 (GCNR) engines. In addition to developing the particular systems, this development provided fundamental information on flow dynamics of pseudo-liquefied powdered fuels and combustion products within the chambers, on feeding said fuels into the combustion chamber and igniting them with decomposition products (in RD-560), along with gaseous flow dynamics (in RD-600). This work formed the basis of the candidate thesis Katorgin defended at the Bauman Moscow State Technical University in 1967 [7].

    The designers encountered a range of principle difficulties. Here is a lis of some of them: [6]

    1. Operation of a gaseous fuel element
    2. Achieving criticality in a gas-core reactor
    3. Achieving stable functioning of the gas-core reactor
    4. Maintaining functionality of components and subsystems at extreme temperatures
    5. Ensuring resistance of construction materials to corrosion
    6. Thermal protection of the nozzle and MHD generator
    7. Separation of fission products in closed-cycle GCNRs

    In 1963-1973 the GCNR and gas-core reactor unit of Energomash included approximately 90 people. That period saw intense experimental and production work on preparing reactor testing that was due to launch in 1975. However, in 1974 Energomash began developing the RD-170/171 – a high-performance kerolox rocket engine for the Energiya-Buran system (along with the Zenit booster, and later for ULA Atlas V in the form of RD-180, for Antares and Angara in the form of RD-190 and for Soyuz-2.1v in the form of RD-193. So much for abandoning cryogenic fuels! – ed.), which caused GCNR research to be halted and the relevant section reduced to 30 people. Over eight years the funding was only sufficient for on-paper studies, resulting in a considerable loss of technological, industrial and experimental groundwork. [5,6]

    Starting in 1982 full-scale development was resumed, and the restored design and development unit spent two years recovering the technological and experimental base. However, in late 1989 funding was cut almost entirely. Neither did any of the programs in the United States reach even the small-scale demonstration experiment stage.

    It was expected that the GCNR would consist of one or several reaction chambers surrounded by a neutron moderator-reflector. The nuclear fuel inside the chambers would be suspended in a plasma state, without contacting the chamber walls, in a quantity sufficient for a self-sustained chain reaction. The reaction mass would flow through the gap between the fissioning plasma and the chamber walls. Reaction mass heating is through radiative energy transfer, with average temperature at chamber egress reaching 104 K. Absorption of radiated heat also provides thermal protection for the chamber walls.

    The key problem in developing the gas-core reactor was minimizing the loss of fission fuel, which had to be kept within tens of percent of reaction mass flow. Acceptable loss level was to be ensured by laminarization of the inbound reaction mass flow, profiling the field of its initial velocities, an external magnetic field, appropriate choice of working materials, and chamber geometry. Loss of fission fuel was to be compensated by its further input in either liquid form (at 1500 K) or as a paste-like powder mix with a NaK eutectic.

    Spaceborne powerplants were designed along both open-cycle and closed-cycle lines. If the working fluid is ejected through a rocket nozzle, then the system is an open-cycle rocket motor. The working fluid if hydrogen that, for the purposes of increased radiative absorption and electric conductivity is doped by NaK and Li vapours along with tungsten powder; this also helps achieve acceptable reactor wall temperature. An NTR of such a design would possess extremely high specific characteristics (Isp on the order of 2000-3000 sec). If the system is designed to eject the hydrogen through a high-efficiency MHD generator, then it is an open-cycle powerplant.

    A closed-cycle powerplant still uses the MHD as a power converter, but all working elements are cycled through isolated loop. In this case we gain a nuclear electric powerplant of impressive efficiency (30-40%) and of low specific mass and working medium expenditure. The additives to the working mass are, among other aspects, intended to improve interaction with the MHD generator. In addition to the reactor and MHD generator, the design inevitable has to include refrigeration, separation and pumping system. The working medium is a mixture of NaK steam and helium. The excess heat is dumped into space via radiators. The power produced can be utilized for a variety of purposes, not the least to power an electric rocket.

    And advantage of gas-core systems over solid-fuel rods in closed-loop powerplants is the ability to considerably extend uninterrupted operation via continuous input of fresh fission fuel to replace the extracted reaction products.

    The conceptual design of a nuclear engine-powerplant system for a manned Mars expedition is the latest-dated, and encompasses all past experience. The design is based around a combined single cavity solid-and-gas-core transforming reactor massing 57.5 t. Gross heat output 2.14 GWt. Solid fuel assemblies, arranged in a ring around the reactor vessel and mounted on an input-extraction system, provide the necessary level of neutron flux for criticality at start-up, before the fuel is introduced to the gaseous fuel element cavity. As the fission fuel is introduced and accumulated in the central cavity, i.e. as a plasma zone appears and the gaseous fuel element is formed, the solid fuel rods are retracted from the active zone, and the reactor becomes a pure gas-core system.

    Thanks to the transforming design the system has two modes of operation:

    • thrust (gas-core) mode developing 17 tf (according to other sources, 600 tf [8]) with an Isp of 2000 sec – for boost and deceleration stages of the trajectory;
    • energetic (solid-core) mode developing 200 kWt of electric power for supplying the needs of a spacecraft with no loss of working medium – for coast stage of the trajectory. This mode involved the operation a closed-cycle gas turbine circuit with a He-Xe mixture as working medium, thermal energy conversion at 20% and radiative cooling via the Brayton cycle.

    In thrust mode, electric energy is generated by a 25 MWt MHD generator integrated into the nozzle, with electrodes and excitement busses oriented down the nozzle’s throat. [6,9]

    RD-600 SCHEMATIC [10] (Figure 1)

    Power block layout: 1 – drive electromotors; 2 – feed-screw; 3 – retractable solid fuel rods; 4 – radiation shield; 5 – coaxial coils; 6 – reaction zone; 7 – reactor structure; 8 – solenoid; 9 – carbon fiber reinforcement coiling; 10 – solenoid heatshield; 11 – lateral moderator-reflector; 12 – high-temperature molybdenum bulkhead; 13 – integrated MHD generator; 14 – supersonic expander nozzle; 15 – front endcap; 16 – fuel rods (graphite with dispersed uranium carbide); 17 – rear endcap; 18 – channels filled with 3He (!? – ed.) (reactor control system actuators); 19 – electrodes for the multipolar Faraday MHD generator

    A layout of a Mars Expeditionary Complex using a block of two above-described gas-core nuclear systems is depicted in Figure 2. At assumed payload of 150 tons for this mission type, approximate mass of the complex in Earth orbit would be 520-540 t depending on launch date. For comparison, an SCNR results in 730-800 t and chemical rockets in 1700-2500 t.

    MEC LAYOUT (Figure 2) (the awful quality is inherited from the original; curiously, it’s annotated in both English and Russian. The scale at the bottom-right is 10 m. Going left-to-right/aft-to-bow, the spacecraft consists of a pair of engines, a V-shaped radiator array around a load-bearing truss that also contains the lithium tanks. Everything between that and the truss section further to the right are hydrogen tanks (jettisonable? There are three distinct sets); the foremost section costs of an orbital habitat with an Earth Return Vehicle stuck to one side and a very vaguely-depicted Mars Landing Vehicle on the other. – ed.)

    Gas-core reactor and derived GCNR development strategy was based around three incremental phases. The initial stage involved the still-operational unique testing facility based around an Impulse Graphite Reactor (IGR) at the Semipalatinsk nuclear range, Kazakhstan. It involved brief (up to 5 sec) operational tests of reduced-scale modes of gaseous fuel elements, up to 100 mm in diameter and 250 mm in length.

    The second phase would involve constructing a new IGR-type reactor called Nephrite, capable of testing samples thrice the size and for an order of magnitude greater durations.

    The final stage would involve a full-scale prototype testbed gas-core or, more precisely, combined solid-and-gas-core reactor dubbed Lampa (before you ask, no other mention of lightbulbs or quartz in anything I’ve found – ed.), with an active zone capable of housing a “dead zone” type, self-contained gas-core fuel element.

    The last two stages would take place at the Baikal-2 stand complex, also at Semipalatinsk. Baikal-2 has had significant research invested into it, with considerable attention paid to safety concerns, primarily radiological and nuclear; in particular, the system was built entirely for closed cycle.

    The preparation for first stage of practical testing of a scale model of a gaseous fuel element in the IGR reactor took the most time and resources. The experimental ampula, containing a model of a fuel elements and all the requisite systems, was to be located into the vertical canal at the centre of the reactor. Over the course of the experiment a displacement-type system was to introduce the fission fuel into the working chamber, located in the middle of IGR’s own active zone. The furl could be a paste composed of small particulate powder of uranium and alkaline metals, or liquid uranium heated before introduction into the chamber. The fuel introduction tract had highly effective and compact neutron shielding in order to prevent overheating of fuel and the surrounding container. Primary dimensions of the interior of the working chamber: diameter 80 mm, length 240 mm. The uranium-containing stream, once injected into the chamber, would be hit by the intense neutron stream, heat up, vaporize and ionize. Radiation from the plasma heated up the working medium. The conical inner wall of the entry section of the working chamber was composed of a high-melting-point alloy. The wall was made permeable to allow injection of hydrogen and helium alongside the fuel. This prevented the formation of a recircularization zone in the fuel vaporization are, and inflow turbulence. The incoming hydrogen, on the other hand, provided a coaxial boundary layer that isolated the chamber walls from the primary uranium plasma stream.

    The cylindrical section of the working chamber had an ablative coating along its interior, providing reliable protection to the outer structure, including in cases of metallic uranium condensing on the ablative material (by the ablation forcing uranium back into the primary stream).

    Upon exiting the chamber the high-temperature stream of reaction mass was to enter the condenser. The walls of the condenser had rows of slots used to inject gaseous hydrogen for the purpose of dilution. Furthermore, the interior of the condenser also had the anti-uranium coating. To reduce heat flow from uranium while passing through the condenser, the exterior of the condenser was also equipped with neutron shielding. The resultant gas mixture containing fission products would then be forced through a transonic nozzle towards the filtration system located near the bottom neutron shield of the reactor. Large particles would be intercepted by the inertial traps, while smaller ones would be caught in metal-ceramic filter cartridges. The use of a transonic nizzle would stabilize the pressure in the working chamber should hydraulic resistance of filtering cartridges change over the course of the experience. Gaseous products would then be routed to the test stand exhaust containment system. To limit head generation and filter heat-up it was also equipped by stationary lateral and bottom neutron shielding.

    Once the fission fuel was exhausted and the test was completed, the shutdown was performed by cooling the heat-producing fuel residue in the ampule filter with a stream of gas.

    The experimental ampule (Figure 3), produced at a pilot plant, had a diameter of 185 mm and a length of 6500 mm, and included the following components: fission fuel feed system, working chamber, condenser, and filter. This, along with communication systems, measurement sensors and overall assembly elements, was packaged into the airtight shell. It was assumed that the requisite supply of the fission fuel would be loaded into the ingress tract of the ampule immediately before commencing the operation. After the test, all solid and liquid products are retained within the filter. Therefore, radiation safety across all stages of operation is ensured through localization and containment of the fission fuel and the bulk of the fission products within the ampule’s interior. Leakage of radioactive substances into the environment was completely excluded.

    The central canal of the IGR impulse reactor, which would hold the experiment ampule, had a water-cooled airtight shell separating it from the uranium-graphite cladding of the active zone. The upper part of the ampule mounted the connections to the test stand’s communications.

    Considerable attention was paid to safety measures preventing damage of IGR reactor core and radioactive contamination of testing facilities in case of possible failure modes of functional subsystems within the experimental ampule.

    Two complete test packages with miniaturized fuel elements were completed and ready for delivery to the test facility (Figure 5) (missing here, see same at – ed.). Specialized test stand equipment kits and expended experimental radioactive material handling and transport equipment were already dispatched there. In addition to the experimental ampule, a draft design of an advanced gas-core fuel element with a “dead space” cooling area and magnetic stabilization was also completed



    1. Б. Е. Черток "Ракеты и люди. Книга 4 Лунная гонка"
    3. Первушин "Битва за звёзды. Космическое противостояние"
    4. Л. Гильберг "Покорение неба", стр. 325-326
    6. "Двигатель", "Газофазные ядерные двигатели для космических аппаратов", 5 1999 г.
    9. "Двигатель", "Газофазные ядерные двигатели для космических аппаратов", 6 1999 г.

    I’ve tracked down the citations. 1 is Boris Chertok’s somewhat less technical autobiographical work covering the Soviet rocket program since he was hunting for leftover V-2s alongside Sergei Korolev; it’s available at 2 and 8 are the same master list of Energomash engines I started from. 3 and 4 are aerospace tech populariser paperbacks. 5 is a dead link to the same site I got Glushko’s correspondence from. 7 and 10 are obsolete links to a periodical that’s behind a paywall; the up-to-date links are and

    6 and 9 are from a journal with the telling name Engines, a two-part article about GCNRs by Grigory Lioznov, an Energomash engineer whom we’ve heard about earlier; they’re available online at and Aside from several more images, I’ve gleaned two very interesting sections not covered above:

    Miniaturization and reduction in mass of GCNRs are facilitated by:

    Use of uranium-233 fission fuel

    •Maximization of use of metallic beryllium, including large-crystal beryllium, in the moderator-reflector assembly, with the rest composed of graphite

    •Maximization of use of metals with improved isotope composition and high melting temperatures in the design of reaction chamber interior, and of high-durability titanium alloys and reinforcing carbon composites in the reactor frame.

    •Use of hyperconductive aluminium (0.9999 purity) in high-current magnetic stabilization, MHD excitation and turbopump power feed systems due to its ability to conduct up to 50-100 A/mm2 when cooled by liquid hydrogen, while developing less than a tenth of resistivity of copper.

    It is obvious that the temperature extremes in operation of many GCNR components and the highly chemically aggressive environment (molten uranium, high-pressure hydrogen, alkali metals) required significant materials science investigations. As a result, high melting point alloys based on tantalum-tungsten-hafnium as well as niobium were developed for the fission fuel feed system. Certain areas of the reactor vessel have necessitated the development of heat-resistant porous materials based on tungsten and molybdenum, for the high-temperature fuel filters – on nickel and nichrome.

    Estimated parameters of a gas-core fuel element
    Pressure in reaction chamber, kgf/cm2200
    Uranium expenditure, g/sec200
    Hydrogen expenditure in the reaction chamber, g/sec10
    Velocity of fuel when entering the reaction chamber, m/s1.7
    Power, kWt1,000
    Share of vaporized uranium in the egress flow, %80
    Temperature of uranium plasma, K8—10×103(unclear – ed.)
    Thermal neutron flow, neutrons/cm2/sec1015(unclear – ed.)

    Final reflections? Just… holy hell. You never know if that old Soviet closet has a skeleton or a suit of Mobile Infantry powered armour in it.

    Yours sincerely,
    Denis Danilov


    These acronyms are odd looking because the original words are in Russian.

    • GFTE: gas-phase fuel rod
    • GMFR: gas-phase nuclear reactor
    • SFSF: structure in high-temperature
    • TFTS: solid-phase fuel assemblies
    • YACEU: radioactive nuclear power and nuclear space power plants
    • YARD: nuclear rocket engine

    The motor power plant of the open circuit (Fig. 1) includes a single-cavity reactor with an annular output channel and a gas-phase fuel rod (GFTE) with a stagnant nuclear zone of nuclear fuel. Zone stabilization is carried out using a powerful external solenoid. The use of such a scheme for environmental reasons is possible only on spacecraft, but not on carriers launched from the Earth.

    In order to provide energy to various consumers, including a solenoid and an electric pump drive, the unit was supposed to use a combination of a nozzle and an MHD generator. In addition to the circuit difference, the YARD and YACEU differ in the degree of use of the gas flow energy in the MHD generator: in the first case, no more than 2% is converted into electricity, and in the second - 30 ... 40%.

    In installations of a closed circuit (Fig. 2), the energy converter is an MHD generator, and all the working components circulate in a circuit that does not have connection with the external environment (so it does not unnecessarily spray fissioning uranium and fission fragments all over tarnation). In this case, we obtain YACEU, which has a very high efficiency (30:40%), low values ​​of the specific gravity of the converter and specific consumption of the working fluid. Additives introduced into the working fluid, among other things, are designed to promote MHD-interaction. In addition to the gas-phase reactor and the MHD generator, refrigerators, separators and pumps must be present in the design. The working fluid is NaK vapor mixed with helium. Released excess heat is discharged into outer space using emitters. The energy produced is used for various purposes, one of its consumers may be an electric rocket engine.

    The advantage of using GFNR in closed circuits, in which gaseous gas is used instead of solid fuel rods, is the fundamental possibility of ensuring a very long-term operation due to appropriate fueling instead of nuclear reaction products removed from the circuit to the external environment

    (Nyrath note:solid fuel rods become clogged with nuclear reaction products (nuclear poisons) when 15% of the fuel has been burnt (nuclear fission). A clogged rod will not support fission. The rod has to be removed and taken to a reprocessing plant. For this reason NASA's reusable nuclear shuttle has an engine life of only 10 Terra-Luna round trips before disposal, even though the rods still contain 85% of the expensive uranium-235 unburnt. Extracting the rods for reprocessing is too dangerous in NASA's eyes, they just send the nuclear shuttle into a graveyard orbit.

    The same limit applies to a nuclear power plant. Ground based plants periodically halt operations when the solid fuel rods become clogged. They then spend a few months carefully opening the reactor, removing the clogged rods, replacing them with fresh rods, then carefully reassembling the reactor. The clogged rods are sent to a reprocessing plant to extract the unburnt U235 and using it to fabricate fresh rods. The rest of the rod goes to a long term nuclear waste disposal site.

    If I am reading the above sentence correctly; they are implying that the fuel gas, after one pass through the MHD generator, can be passed through an on-board refinery. The product is already gaseous, as opposed to solid fuel rods, which simplifies refining. In the refinery the nuclear reaction products clogging the gas can be filtered out, and the unburnt U235 can be sent on another pass through the MHD generator. The alternative is removing the gas "from the circuit to the external environment", i.e., wastefully jettisoning the gas into deep space along with all its expensive unburnt U235.

    Now that my nose has been rubbed in the fact, I realize that a nuclear lightbulb gas core engine should also require an on-board reprocessing plant. But in all the nuclear lightbulb documents I've read, a re-processor is conspicuous by its absence. I'm going to have to review the documents.)

    The fact that in closed circuits the requirement for the removal of nuclear fuel from the reactor together with the working fluid is less stringent than in the open is also significant. This allows us to consider a more simple organization of processes that allow a greater degree of mixing of nuclear fuel and working fluid. In this case, there is no need for magnetic stabilization — the plasma zone from the stagnant turns into a jet zone. The use of several such zones (multi-cavity reactor) improves the overall size characteristics of the SFSF.

    It is known that there is a definite relationship between the thermal power of the reactor and the possibilities of providing an acceptable temperature condition for the structural elements. Research has found that the optimal thermal capacity of an open-circuit GFNR should be no less than 2 GW, and a closed one — 300 MW (with a pressure in the working chamber of about 1000 kgf / cm2).

    The conceptual development of a nuclear propulsion system to support the Martian expedition is the latest in time, incorporating all previous experience. The installation is based on a combined single-cavity gas-phase-solid-phase reactor of a transformable structure weighing 57.5 tons (Fig. 3). Thermal power of the reactor is 2.14 GW. Solid-phase fuel assemblies (TFTS), placed in a ring around the central cavity of the reactor and equipped with drive mechanisms, provide the necessary level of neutron flux and criticality at start-up when there is no nuclear fuel in the cavity of a gas-phase fuel element. With the supply and accumulation in the central cavity of a nuclear fuel, i.e. the formation of the plasma zone and the formation of a gas-phase fuel element, TFTS from the active zone are extracted.

    Thanks to the transformable design, the installation can operate in two modes:

    • Thrust Mode: a 17 tonne-force motor (gas phase) with a specific impulse of 2000 seconds
    • Coast Mode: energy (solid-phase) with an electric power of 200 kWe to meet the internal needs of the spacecraft without spending the working fluid. This mode is provided by a closed gas turbine circuit with a helium-xenon mixture as a working fluid, converting thermal energy into electricity with efficiency % and the discharge of excess heat through the heat radiator (Brighton cycle)

    In thrust mode of operation, the power supply is provided by a 25 MW multi-pole MHD generator built into the nozzle with electrodes and field buses oriented along the nozzle-forming elements.

    Nuclear Salt Water
    20% UTB
    Exhaust Velocity66,000 m/s
    Specific Impulse6,728 s
    Thrust12,900,000 N
    Thrust Power425.7 GW
    Mass Flow195 kg/s
    Total Engine Mass33,000 kg
    Uranium Tetrabromide
    ReactorGas Core
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorPusher Plate
    Specific Power0.08 kg/MW
    90% UTB
    Exhaust Velocity4,700,000 m/s
    Specific Impulse479,103 s
    Thrust13,000,000 N
    Thrust Power30.6 TW
    Mass Flow3 kg/s

    This concept by Dr. Zubrin is considered far-fetched by many scientists. The fuel is a 2% solution of 20% enriched Uranium Tetrabromide in water. A Plutonium salt can also be used.

    Just to make things clear, there are two percentages here. The fuel is a 2% solution of uranium tetrabromide and water. That is, 2 molecules of uranium tetrabromide per 100 molecules of water.

    But the uranium tetrabromide can be 20% enriched. This means that out of every 100 atoms of uranium (or molecules of uranium tetrabromide), 20 are fissionable Uranium-235 and 80 are non-fissionable uranium. If it is 90% enriched, then 90 atoms are Uranium-235 and 10 atoms are non-fissionable. As a side note, 90% enriched is considered "weapons-grade".

    The fuel tanks are a bundle of pipes coated with a layer of boron carbide neutron damper. The damper prevents a chain reaction. The fuel is injected into a long cylindrical plenum pipe of large diameter, which terminates in a rocket nozzle. Free of the neutron damper, a critical mass of uranium soon develops. The energy release vaporizes the water, and the blast of steam carries the still reacting uranium out the nozzle.

    It is basically a continuously detonating Orion type drive with water as propellant. Although Zubrin puts it like this:

    As the solution continues to pour into the plenum from the borated storage pipes, a steady-state conditions of a moving detonating fluid can be set up within the plenum.

    He also notes that it is preferable to subject your spacecraft to a steady acceleration (as with the NSWR) as opposed to a series of hammer-blow accelerations (as with Orion).

    The controversy is over how to contain such a nuclear explosion. Zubrin maintains that skillful injection of the fuel can force the reaction to occur outside the reaction chamber. He says that the neutron flux is concentrated on the downstream end due to neutron convection. Other scientists are skeptical.

    Naturally in such a spacecraft, damage to the fuel tanks can have unfortunate results (say, damage caused by hostile weapons fire). Breach the fuel tubes and you'll have a runaway nuclear chain reaction on your hands. Inside your ship.

    The advantage of NSWR is that this is the only known propulsion system that combines high exhaust velocity with high thrust (in other words, it is a Torchship). The disadvantage is that it combines many of the worst problems of the Orion and Gas Core systems. For starters, using it for take-offs will leave a large crater that will glow blue for several hundred million years, as will everything downwind in the fallout area.

    Zubrin calculates that the 20% enriched uranium tetrabromide will produce a specific impulse of about 7000 seconds (69,000 m/s exhaust velocity), which is comparable to an ion drive. However, the NSWR is not thrust limited like the ion drive. Since the NSWR vents most of the waste heat out the exhaust nozzle, it can theoretically produce jet power ratings in the thousands of megawatts (meaning it is not power limited, like other nuclear propulsion). Also unlike the ion drive, the engine is relatively lightweight, with no massive power plant required.

    Zubrin suggests that a layer of pure water be injected into the plenum to form a moving neutron reflector and to protect the plenum walls and exhaust nozzle from the heat. One wonders how much protection this will offer.

    Zubrin gives a sample NSWR configuration. It uses as fuel/propellant a 2% (by number) uranium bromide aqueous solution. The uranium is enriched to 20% U235. This implies that B2 = 0.6136 cm-2 (the material buckling, equal to vΣfa)/D) and D = 0.2433 cm (diffusion coefficent).

    Radius of the reaction plenum is set to 3.075 centimeters. this implies that A2 = 0.6117 cm-2 and L2 = 0.0019. Since exponential detonation is desired, k2 = 2L2 = 0.0038 cm-2. Then k = U / 2D = 0.026 cm-1 and U = 0.03.

    If the velocity of a thermal neutron is 2200 m/s, this implies that the fluid velocity needs to be 66 m/s. This is only about 4.7% the sound speed of room temperature water so it should be easy to spray the fuel into the plenum chamber at this velocity.

    The total rate of mass flow through the plenum chamber is about 196 kg/s.

    Complete fission of the U235 would yield about 3.4 x 1012 J/kg. Zubrin assumes a yield of 0.1% (0.2% at the center of the propellant column down to zero at the edge), which would not affect the material buckling during the burn. This gives an energy content of 3.4 x 109 J/kg.

    Assume a nozzle efficiency of 0.8, and the result is an exhaust velocity of 66,000 m/s or a specific impules of 6,7300 seconds. The total jet power is 427 gigawatts. The thrust is 12.9 meganewtons. The thrust-to-weight ratio will be about 40, which implies an engine mass of about 33 metric tons.

    For exponetial detonation, kz has to be about 4 at the plenum exit. Since k = 0.062 cm-1, the plenum will have to be 65 cm long. The plenum will be 65 cm long with a 3.075 cm radius, plus an exhaust nozzle.

    Zubrin then goes on to speculate about a more advanced version of the NSWR, suitable for insterstellar travel. Say that the 2% uranium bromide solution used uranium enriched to 90% U235 instead of only 20%. Assume that the fission yield was 90% instead of 0.1%. And assume a nozzle efficency of 0.9 instead of 0.8.

    That would result in an exhaust velocity of a whopping 4,725,000 m/s (about 1.575% c, a specific impulse of 482,140 seconds). In a ship with a mass ratio of 10, it would have a delta V of 3.63% c. Now you're talkin...

    Ken Burnside: In my game universe, the engineers call the pumps that feed Uranium Tetrabromide solution into the reaction chamber "Wileys", reputedly after the engineer who first made them safe to use and maintain.

    More than likely, it's after the coyote of the same name...

    Winchell Chung: An appropriate name for what are basically atomic squirt-guns.

    From a thread in SFConSim-l (2002)
    Zubrin NSWR
    Zubrin NSWR
    Exhaust Velocity78,480 m/s
    Specific Impulse8,000 s
    Thrust8,696,900 N
    Thrust Power0.3 TW
    Mass Flow111 kg/s
    Total Engine Mass495,467 kg
    Frozen Flow eff.80%
    Total eff.80%
    Uranium Tetrabromide
    ReactorGas Core
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorPusher Plate
    Specific Power1.45 kg/MW

    The illustration shows the vision of Robert Zubrin: a rocket riding on a continuous controlled nuclear explosion just aft of a nozzle/reaction chamber.

    The propellant is water, containing dissolved salts of fissile uranium or plutonium. These fuel-salts are stored in a tank made from capillary tubes of boron carbide, a strong structural material that strongly absorbs thermal neutrons, preventing the fission chain reaction that would otherwise occur.

    To start the engine, the salt-water is pumped from the fuel tank into an absorber-free cylindrical nozzle. The salt-water velocity is adjusted as it exits the tank so that the thermal neutron flux peaks sharply in the water-cooled nozzle.

    At critical mass (around 50 kg of salt water), the continuous nuclear explosion produces 427 GWth, obtaining a thrust of 8600 kN and a specific impulse of 8 ksec at a thermal efficiency of 99.8% (with open-cycle cooling). Overall efficiency is 80%.

    Robert Zubrin, "Nuclear Salt Water Rockets: High Thrust at 10,000 sec ISP," Journal of the British Interplanetary Society 44, 1991.

    You need much more propellant than fuel, 22,000 times more in the case of the Zubrin without open cycle cooling, and 44,000 times more if open cycle cooling is used.

    The Zubrin drive exhaust (without open cycle cooling) contains 108 kg/sec of water, but only about 5 grams/sec of uranium.

    (This is from a quick calculation: mass flow equals the Zubrin thrust (8.7 meganewtons) divided by the exit velocity (80 km/sec) = 108 kg/sec. But the fissioning energy can be estimated from the Zubrin total power of 427 GW divided by the energy content of Uranium 235 of 83 TJ/kg.)

    Dr. Zubrin responded, and he defends the performance of the Zubrin drive as depicted in the game (as high thrust & high specific impulse rocket with low mass and low radiators).

    1). In U235 fission, only about 2% of the energy goes into neutrons (unlike D-T fusion).

    2). The design uses a pusher plate or open nozzle, like an Orion drive. Or magnetic confinement (since most of the energy is released as a plasma). Therefore, the opportunity to absorb heat is low.

    3) Many of the neutrons that are intercepted would sail through the pusher plate, rather than be absorbed as waste heat.

    4) No lithium should be in the outer water, because this would poison the fission reactions.

    5). Because the design does not use a heat engine cycle, the radiators could be far hotter than ones in the game. He suggested graphite at 2500 K°. That would drop the required radiating area by a factor of 40 (2.5 to the fourth power), which means that the radiator could be the first wall itself.

    Dr. Zubrin went on to say the chief disadvantage is the expense of the fuel (like He3-D and antimatter drives).

    Philip Eklund, from a discussion on the High Frontier Yahoo group about the NSWR drive in High Frontier

         "So anyway, we were passing through the outer Kirkwood Gap, totally the a** end of nowhere. I'm trying to catch some rack time, XO has the conn, nice boring trip to Europa." The CO of U.S.N.A.S. Saskatchewan tipped back another shot of Scotch and continued his story. "Totally routine, right? No problems at all. So then, all of the sudden, the whole ship gets racked. Meteoroid. Big one, too, maybe a centimeter across."

         The captains seated around the table, two Americans including Fitzthomas, an Indian, three Chinese, and the South African, all clucked and groaned.

         "Well, we got lucky and it missed the crew compartment, but by the time I get to command the chief engineer is screaming over the intercom that it holed tank one, busted three tubes, and we've got nuke juice pooling and we have to dump the tank. Problem is, we're running at top speed and if we dump the tank, we don't have enough propellant to stop at Europa. We'd have to ride all the way out to Neptune, sling around, and hope someone from the inner solar system has dispatched a tanker to intercept us on the return trip, and we don't have near enough consumables for that."

         "So what did you do, mate?" said the South African.

         "I told the chief he had to fix the tank or we'd all starve before we could stop the damn ship. Well, he screams some more that we don't have time, and I tell him his choice is fix the tank or die real fast in a runaway, because we're not going to die slow in the void. So he grabs a crew, stuffs them into suits, and crawls out onto the tank. They punch some holes in it to let the juice drain instead of pool, but it's still leaking like a f***** and the water's evaporating and leaving uranium crusted all over everything. So he radios command and says, 'It's still leaking, and all this uranium crud is going to accumulate into a critical mass somewhere, so we still have to drop the tank.' Meteoroid busted open three valves, you see. No way to stop the leak. And I tell him again, that's no good, and by now astrogation has confirmed it and the XO has tallied up the consumables and I know for sure we don't have enough for an unscheduled trip to Neptune.

         "So he says something about how he's not a miracle worker, and I tell him he damn well has to be. Lo and behold, he and his crew go ahead and do something crazy and it works."

         "What was that?" said the South African.

         "They take torches to the tank. The plug up the broken pipes as best they can, and then they go ahead and cut away the smashed cells. Just cut it off and jettison it into space, and suddenly the propellant that's still leaking is leaking right into space. We have lousy flow through the tank and the braking burn is going to be real tricky, but we can make Europa. I put the chief up for a commendation medal for figuring that out on the fly and saving our asses."

         The other captains nodded their approval at the chief's quick thinking. Good chiefs prevented accidents; great ones prevented disasters.

         "Is the chief's name Mr. Scott, by any chance?" said one of the Chinese captains.

         Commander George Allen, New Jersey's full blooded Cherokee XO, drifted into the command deck from astrogation, where he'd been monitoring the final approach to Hektor. He took his place at the copilot station and put on his headset. Fitzthomas toggled his direct channel to Allen's station.

         "How was the approach?"
         "We wasted too much propellant before the chain reaction started. I think Pennai should inspect the nozzles and pumps before we get underway again."
         "What does Pennai say about it?"
         Pause. "Pennai thinks the fuel is dirty."
         "Is it?"
         "It was certified 90% enriched at Roosevelt Station."
         "Is there any way to test it here?"
         "No sir. Not without a centrifuge."
         "How does Pennai know, then?"
         "Some engineering technobabble about neutron flux and reaction rate. I couldn't follow a tenth of what she said."
         Fitzthomas considered that. "Have her inspect the pumps and nozzle alignment. If they pass, then we might have a fuel problem."

         "Captain," said Allen, "Thought you'd like to know: Pennai just inspected the entire fuel line. Everything there is in order."
         "So what are you telling me, George?"
         "I think we have dirty fuel."
         "What's her recommendation?"
         "She wants to drain the tanks and top up with the good stuff. But I can't—"
         "Write that request, I know. The CO has to. Where's Pennai now?"
         "Racked out. She has the midwatch tonight."
         "After her watch tonight, she has four days of leave."
         "Sir, she's supposed to be OOW all day Wednesday."
         "I'll take that shift. She was right, we were wrong. She deserves to be rewarded. When I get back I'll write up a request and have it to the fuelmaster by tomorrow AM."

         "Do you have to return to your ship?"
         "Yeah. Dirty fuel, God damn it. Wait until I get my hands on the fuelmaster at Roosevelt."

         (Admiral Castro said) "Anyway, I saw your chief engineer's report. I passed it back to Fleet. The fuelmaster at Roosevelt Station is going to have a lousy day tomorrow. There's also a bulletin going out to the entire fleet. Everyone who tanked up at Roosevelt near the same time you did should keep a close eye on his reaction rate. Your Lieutenant Pennai might be up for a commendation letter in her file."

         Duvalier left Ortiz main engineering and vaulted down the access tube to the reactor room. The tube ran down the ship's spine, surrounded by megaliters of water enriched with uranium salts in highly complex tanks made of neutron absorbing material. In his head, he knew the tube was the safest part of the ship, shielded from the worst the universe could throw at it by dozens of meters of water. In his head, he knew the fuel, so long as it didn't pool into a critical mass somewhere in the thousands of kilomters of pipes on all sides of him, emitted only low intensity alpha rays which couldn't penetrate his own skin, let alone the aluminum skin of the pressure tube. It was all perfectly safe, so far as anything in space could be safe. He knew that in his head.

         His b***s, however, hadn't gotten the memo. His testicles tried to crawl up into his body every time he climbed through the hatch.

    From The Last Great War by Matthew Lineberger (not yet published)

    Fission Fragment

    Fission Fragment

    George Chapline
    Specific Impulse1,000,000 sec
    Exhaust velocity9,810,000 m/s

    This is from Fission Fragment Rockets — A Potential Breakthrough

    All of the other nuclear thermal rockets generate heat with nuclear fission, then transfer the heat to a working fluid which becomes the reaction mass. The transfer is always going to be plagued by inefficiency, thanks to the second law of thermodynamics. What if you could eliminate the middleman, and use the fission heat directly with no transfer?

    That what the fission fragment rocket does. It uses the hot split atoms as reaction mass. The down side is that due to the low mass flow, the thrust is minuscule. But the up side is that the exhaust velocity is 3% the speed of light! 9,810 kilometers per second, that's like a bat out of hell. With that much exhaust velocity, you could actually have a rocket where less than 50% of the total mass is propellant (i.e., a mass ratio below 2.0).

    The fission fragment is one of the few propulsion systems where the reaction mass has a higher thermal energy than the fuel elements. The other notable example being the Pulsed NTR.

    Dr. Chapline's design use thin carbon filaments coated with fission fuel (coating is about 2 micrometers thick). The filaments radiated out from a central hub, looking like a fuzzy vinyl LP record. These revolving disks were spun at high speed (1 km/sec) through a reactor core, where atoms of nuclear fuel would undergo fission. The fission fragments would be directed by magnetic fields into an exhaust beam.

    The drawback of this design is that too many of the fragments fail to escape the fuel coat (which adds no thrust but does heat up the coat) and too many hit the carbon filaments (which adds no thrust but does heat up the filaments). This is why the disks spin at high speed, otherwise they'd melt.

    Dusty Plasma
    Thrust22 N
    Thrust Power0.2 GW
    Mass Flow1.00e-06 kg/s
    Specific Power55 kg/MW
    Thrust344 N
    Thrust Power2.6 GW
    Mass Flow2.30e-05 kg/s
    Specific Power3 kg/MW
    Exhaust Velocity15,000,000 m/s
    Specific Impulse1,529,052 s
    Total Engine Mass9,000 kg
    Uranium 235
    ReactorGas Core
    MHD Choke
    Remass AccelFission-Fragment
    Thrust DirectorMagnetic Nozzle

    Rodney Clark and Robert Sheldon solve the problem with their Dusty plasma bed reactor (report).

    You take the fission fuel and grind it into dust grains with an average size of 100 nanometers (that is, about 1/20th the thickness of the fuel coating in dr. Chapline's design). This does two things [A] most of the fragments escape and [B] the dust particles have such a high surface to volume ratio that heat (caused by fragments which fail to escape) readily dissipates, preventing the dust particles from melting.

    The dust is suspended in the center of a reaction chamber whose walls are composed of a nuclear moderator. Power reactors will use beryllium oxide (BeO) as a moderator, but that is a bit massive for a spacecraft. The ship will probably use lithium hydride (LiH) for a moderator instead, since is only has one-quarter the mass. Probably about six metric tons worth. The dust is suspended electrostatically or magnetically by a containment field generator. The dust is heated up by radio frequency (RF) induction coils. The containment field generator will require superconductors, which will probably require a coolant system of its own.

    The dust particles are slow and are relatively massive, while the fission fragments are fast and not very massive at all. So the magnetic field can be tailored so it holds the dust but allows the fission fragments to escape. Magnetic mirrors ensure that fragments headed the wrong way are re-directed to the exhaust port.

    One valuable trick is that you can use the same unit for thrust or to generate electricity. Configure the magnetic field so that the fragments escape "downward" through the exhaust port and you have thrust. Flip a switch to change the magnetic field so that the fragments escape upward into deceleration and ion collection electrodes and you generate electricity. As a matter of fact, it is so efficient at generating electricity that researchers are busy trying to adapt this for ground based power plants. But I digress.

    The dust is only sufficient for a short period of critical nuclear reaction so it must be continuously replenished. The thermal energy released by fission events plus heat from collisions between fission fragments and dust grains create intense heat within the dust cloud. Since there is no core cooling flow, the reactor power is limited to the temperature at which the dust can radiatively cool itself without vaporizing. The interior of the reaction chamber walls will protected by a mirrored (95% reflection) heat shield attached to a heat radiator. The outer moderator layer will have its own heat shield.

    Clark and Sheldon roughed out a propulsion system. It had six tons for the moderator, 2 tons for radiators and liquid metal cooling, 1 ton for magnets, power recovery, and coils, for a grand total of 9 tons. The reaction chamber will be about 1 meter in diameter and 10 meters long. The moderator blanket around the chamber will be about 40 centimeters thick. The thrust is a function the size of the cloud of fissioning dust, and is directly related to the power level of the reactor. There is a limit to the maximum allowed power level, set by the coolant system of the reaction chamber. Clark and Sheldon estimate that only about 46% of the fission fragments provide thrust while the rest are wasted. See the report for details.

    In the table, the 550AU engine is for a ten year journey to the Solar gravitational lensing point at 550 astronomical units (so you can use the sun as a giant telescope lens). The 0.5LY engine is for a thirty year trip to the Oort cloud of comets. These are constant acceleration brachistochrone trajectories, the 550AU mission will need a reactor power level of 350 MW and the 0.5LY mission will need 5.6 GW. Don't forget that the engine power is only 46% efficient, that's why the table thrust values are lower.

    Werka FFRE
    First Generation
    Exhaust Velocity5,170,000 m/s
    Specific Impulse527,013 s
    Thrust43 N
    Thrust Power0.1 GW
    Mass Flow8.00e-06 kg/s
    Total Engine Mass113,400 kg
    Plutonium 239
    ReactorGas Core
    MHD Choke
    Remass AccelFission-Fragment
    Thrust DirectorMagnetic Nozzle
    Specific Power1,020 kg/MW
    Propulsion SystemWerka FFRE
    Wet Mass303,000 kg
    Dry Mass295,000 kg
    Mass Ratio1.03
    ΔV138,336 m/s

    Robert Werka has a more modest and realistic design for his fission fragment rocket engine (FFRE). He figures that a practical design will have an exhaust velocity of about 5,200,000 m/s instead of his estimated theoretical maximum of 15,000,000 m/s. His lower estimate is still around 1.7% the speed of light so we are still talking about sub 2.0 mass ratios. Collisions between fission fragments and the dust particles is responsible for the reduction in exhaust velocity.

    Incidentally the near relativistic exhaust velocity reduces radioactive contamination of the solar system. The particles are traveling well above the solar escape velocity (actually they are even faster than the galactic escape velocity) so all the radioactive exhaust goes shooting out of the solar system at 0.017c.

    The dusty fuel is nanometer sized particles of slightly critical plutonium carbide, suspended and contained in an electric field. A moderator of deuterated polyethylene reflects enough neutrons to keep the plutonium critical, while control rods adjust the reaction levels. The moderator is protected from reaction chamber heat by a heat shield, an inner layer composed of carbon-carbon to reflect infrared radiation back into the core. The heat shield coolant passes through a Brayton cycle power generator to create some electricty, then the coolant is sent to the heat radiator.

    The details of Werka's initial generation FFRE can be found in the diagram below. The reaction chamber is about 5.4 meters in diameter by 2.8 meters long. The magnetic nozzle brings the length to 11.5 meters. The fuel is uranium dioxide dust which melts at 3000 K, allowing a reactor power of 1.0 GW. It consume about 29 grams of uranium dioxide dust per hour (not per second). Of the 1.0 GW of reactor power, about 0.7 GW of that is dumped as waste heat through the very large radiators required.

    The second most massive component is the magnetic mirror at the "top" of the reaction chamber. Its purpose is to reflect the fission fragments going the wrong way so they turn around and travel out the exhaust nozzle. Surrounding the "sides" of the reaction chamber is the collimating magnet which directs any remaining wrong-way fragments towards the exhaust nozzle. The exhaust beam would cause near-instantaneous erosion of any material object (since it is electrically charged, relativistic, radioactive grit). It is kept in bounds and electrically neutralized by the magnetic nozzle cage.

    Afterburner Fission Fragment

    Engine Mass
    107,000 kg
    Engine Mass
    (mod oil)
    91,000 kg
    Engine Mass
    268,961 kg
    Reactor Power2.5 GW
    Thrust4,651 N
    Thrust Power730 MW
    32,000 sec
    313,900 m/s
    Mass Flow
    3.12×10-5 kg/s
    Mass Flow
    0.0179 kg/s
    Mass Flow
    0.018 kg/s

    Robert Werka has apparently figured out a new configuration for his fission-fragment rocket engine (FFRE). The report is here.

    As with most engines that have high specific impulse and exhaust velocity, the thrust of a FFRE is pathetically small. Ah, but there is a standard way of dealing with this problem: shifting gears. What you do is inject cold propellant into the exhaust ("afterburner"). The fission fragment exhaust loses energy while the cold propellant gains energy. The combined exhaust velocity of the fission fragment + propellant energy is lower than the original pure fission fragment, so the specific impulse goes down. However the propellant mass flow goes up since the combined exhaust has more mass than the original pure fission fragment. So the thrust goes up.

    Now you have an Afterburner fission-fragment rocket engine (AFFRE).

    As you are probably tired of hearing, this means the engine has shifted gears by trading specific impulse for thrust.

    Shifting Gears
    FFRE527,000 sec43 Newtons
    AFFRE32,000 sec4,651 Newtons

    The heart of the engine is a standard "dusty plasma" fission fragment engine. A cloud of nanoparticle-sized fission fuel is held in an electrostatic field inside a neutron moderator. Atoms in the particles are fissioning like crazy, spewing high velocity fission products in all directions. These become the exhaust, directed by a magnetic nozzle.

    The AFFRE alters this a bit. Instead of a cylindrical reactor core it uses half a torus. Each end of the torus has its own magnetic nozzle. But the biggest difference is that cold hydrogen propellant is injected into the flow of fission fragments as an afterburner, in order to shift gears.

    In the diagram above, the magnetic nozzles are the two frameworks perched on top of the reactor core. It is a converging-diverging (C-D) magnetic nozzle composed of a series of four beryllium magnetic rings (colored gold in the diagram). Note how each frame holding the beryllium rings is shaped like an elongated hour-glass, that is the converting-diverging part. The fission fragment plume emerges from the reactor core, is squeezed (converges) down until it reaches the midpoint of the magnetic nozzle, then expands (diverges) as it approaches the end of the nozzle. At the midpoint is the afterburner, where the cold hydrogen propellant is injected.

    The semi-torus has a major and minor radius of 3 meters. The overall length of the engine is 13 meters. The reactor uses 91 metric tons of hydrocarbon oil as a moderator. This means the heavy lift vehicle can launch the engine "dry" with no oil moderator. In orbit the oil moderator can be easily injected into the reactor, at least easier than building the blasted thing in free fall out of graphite bricks.

    Fission Sail

    Fission Sail

    Antimatter-Driven Sail

    The sail is made of graphite and carbon-carbon fiber, infused with a tiny amount of uranium. It is subjected to a misting of antiprotons. These induce uranium atoms to fission, with the recoil pushing the sail. Since this is nuclear powered, the sail does not have to be kilometers in diameter, five meters will do. 30 miligrams of antiprotons could push the sail to the Kuiper Belt.


    This is from LASL nuclear rocket propulsion program (1956) and Propulsion Systems for Space Flight by William R. Corliss (1960). It is called a "consumable nuclear rocket", a "Fizzer", a "Fizzing Bomb", or a "Burning Wall" rocket.

    This is totally insane. Thank the stars it was never developed. This is sort of a mash-up of a solid-core NTR, a gas-core NTR, a chemical solid rocket, and atomic Primacord. Think of it as a giant nuclear-powered sparkler from hell. It is from those innocent days when the rocket designers wouldn't recognize a bad idea even if it they tripped over it.

    Note that the propellant is lithium hydride, presumably a convenient way to hold hyrogen since cryogenic tanks need refrigeration equipment and all sorts of extra stuff.

    The propellant is NOT lithium deuteride. There is another name for a fission reaction next to a slab of lithium deuteride, it is "thermonuclear weapon." Lithium hydride creates thrust. Lithium deuteride will just go off like an H-bomb, vaporizing the payload and anything else nearby.

    FIZZER 1

    There are three basic types of nuclear rockets. They may be classified by the manner in which the nuclear reaction is contained (Table 5-1).

    The so—called “heat-transfer nuclear rocket” is no more than a nuclear-reactor core through which propellant is passed and heated. Figure 5-14 illustrates this type schematically. The second type is the consumable nuclear rocket. Here the nuclear fissioning occurs directly in the working fluid. Both fission fuel and fission products are released in the exhaust. This type is conceptually midway between the heat transfer reactor and an atomic bomb.

    An example of the solid type is shown in Fig. 5-21. The U—235 fuel is concentrated in a rod running the length of the rocket. A good neutron absorber, such as cadmium, surrounds the fuel. Around this there is a thick layer of propellant, perhaps lithium hydride. To fire the rocket, a section of the cadmium sheath is pulled off the bottom of the fuel cylinder causing the neutron chain reaction to begin. In the region where cadmium is absent, the nuclear reaction will flash the solid materials into a high temperature vapor. The anisotropic expansion of the vapor will produce a forward thrust on the rocket. The average molecular weight of the vapor will be dictated by the relative proportions of uranium and propellant in a given cross section. Usually the ratio of masses will be about 100 kg of propellant for each kilogram of fuel. Once the reaction begins, it will proceed up the body of the rocket as the neutron temperature rises high enough to make the cadmium an ineffectual neutron absorber. Unfortunately, no way has yet been found to prevent the reaction from traveling at velocities exceeding 100 m/sec. The consequence of this fact is that inordinately long rockets would be needed to obtain reasonable accelerations and burn times. For manned systems, it is desirable to keep accelerations below 10 g. Burn times between 10 and 100 sec are satisfactory for many launching missions. If this fundamental problem of slowing down the reaction can be solved, the solid, consumable nuclear rocket is simple and attractive in operation.

    From PROPULSION SYSTEMS FOR SPACE FLIGHT by William R. Corliss (1960)
    FIZZER 2

    Another class of rockets is generally typified by the fizzling bomb concept. At a sufficiently high rate of reaction, a substantial energy release can be achieved by an explosive fission reaction in which the recoiling mass of the reactor itself furnishes the impulse to drive the vehicle. Highly moderated reactors should be used, both for low critical mass and to provide inexpensive material to be ejected. Possible methods lunge all of the way from a single shot, through multiple explosions, to a continuous reaction analogous to a solid propellant chemical rocket. All such schemes are characterized by the incompatibility of reasonable accelerations and economical use of active materiel. Roughly speaking, the time of an explosive nuclear reaction is equal to the shock wave transit time across the reacting zone. Unless some ingenious cushion is built in, this time also characterizes the impulse given to the missile. If this impulse is to give the missile a velocity increment of several thousand feet per second, the resultant accelerations are fairly fantastic.

    In principle, a slower reaction with reasonable fissionable material economy could be achieved with a gaseous reactor that retains preferentially the fissionable fuel, but no schemes yet proposed seem workable.

    Thermal and fast neutron reactors, separated fission products, direct use of fission fragments to heat gas, self-heating mixtures of moderator and fuel (both thrown away), thrust from fission fragment momentum, alpha particle recoil, and electron and ion accelerators were considered for vehicle propulsion systems.

    The LASL, UCRL and Lockheed discussions generally pertained to the use of unconventional means for the production of thrust, such as fissioning gaseous systems, burning "cigarettes" or internally burning reactors, or the use of radiation from bomb bursts to heat a working fluid (BATO) . Martin proposed to obtain thrust from thermonuclear processes achieved in a transient (pulsed) shock wave heated system. In general, none of these imconventional methods offered any real hope of successful achievement in the near future.

    UCRL reported it has concluded the fizzling and gaseous reactors seem to have one order of magnitude of impossibility inherent in their makeup and that they therefore do not appear promising for application to rocket use.


    These are propulsion systems where the thrust is not delivered as a constant burn, but instead in a series of intermittent pulses.

    Note that most pulse propulsion systems can "throttle" their effective thrust by varying the pulse rate from 0 to max.


    Fission Orion
    Exhaust Velocity43,000 m/s
    Specific Impulse4,383 s
    Thrust263,000 N
    Thrust Power5.7 GW
    Mass Flow6 kg/s
    Total Engine Mass200,000 kg
    Uranium 235
    ReactorPulse Unit
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorPusher Plate
    Specific Power35 kg/MW
    Fusion Orion
    Exhaust Velocity73,000 m/s
    Specific Impulse7,441 s
    Thrust292,000 N
    Thrust Power10.7 GW
    Mass Flow4 kg/s
    Total Engine Mass200,000 kg
    FuelD-D Fusion
    Specific Power19 kg/MW
    1959 Orion 1st Gen
    Thrust Power1,600 GW
    Exhaust velocity39,000 m/s
    Thrust80,000,000 n
    Engine mass1,700 tonne
    T/W >1.0yes
    1959 Orion 2nd Gen
    Thrust Power24,000 GW
    Exhaust velocity120,000 m/s
    Thrust400,000,000 n
    Engine mass3,250 tonne
    T/W >1.0yes
    ORION USAF 10m [*]
    Exhaust Velocity32,900 m/s
    Specific Impulse3,354 s
    Thrust2,000,000 N
    Thrust Power32.9 GW
    Mass Flow61 kg/s
    Total Engine Mass107,900 kg
    Wet Mass475,235 kg
    Dry Mass180,975 kg
    Mass Ratio2.63 m/s
    ΔV31,763 m/s
    Specific Power3 kg/MW
    ORION 4K ton battleship
    Exhaust Velocity39,000 m/s
    Specific Impulse3,976 s
    Thrust80,000,000 N
    Thrust Power1.6 TW
    Mass Flow2,051 kg/s
    Total Engine Mass1,700,000 kg
    Specific Power1.09 kg/MW
    ΔV 10 km/s
    Wet Mass4,000,000 kg
    Dry Mass3,100,000 kg
    Mass Ratio1.29 m/s
    ΔV9,941 m/s
    ΔV 21 km/s
    Wet Mass4,000,000 kg
    Dry Mass2,353,000 kg
    Mass Ratio1.70 m/s
    ΔV20,694 m/s
    ΔV 30 km/s
    Wet Mass4,000,000 kg
    Dry Mass1,852,000 kg
    Mass Ratio2.16 m/s
    ΔV30,031 m/s
    ORION 10k ton adv
    Exhaust Velocity120,000 m/s
    Specific Impulse12,232 s
    Thrust400,000,000 N
    Thrust Power24.0 TW
    Mass Flow3,333 kg/s
    Total Engine Mass3,250,000 kg
    Specific Power0.14 kg/MW
    ΔV 10 km/s
    Wet Mass10,000,000 kg
    Dry Mass9,199,000 kg
    Mass Ratio1.09 m/s
    ΔV10,019 m/s
    ΔV 15.5 km/s
    Wet Mass10,000,000 kg
    Dry Mass8,772,000 kg
    Mass Ratio1.14 m/s
    ΔV15,722 m/s
    ΔV 20 km/s
    Wet Mass10,000,000 kg
    Dry Mass8,403,000 kg
    Mass Ratio1.19 m/s
    ΔV20,880 m/s
    ΔV 30 km/s
    Wet Mass10,000,000 kg
    Dry Mass7,813,000 kg
    Mass Ratio1.28 m/s
    ΔV29,616 m/s
    ΔV 100 km/s
    Wet Mass10,000,000 kg
    Dry Mass4,348,000 kg
    Mass Ratio2.30 m/s
    ΔV99,944 m/s
    Orion MAX
    Exhaust Velocity9,800,000 m/s
    Specific Impulse998,981 s
    Thrust8,000,000 N
    Thrust Power39.2 TW
    Mass Flow0.82 kg/s
    Total Engine Mass8,000 kg
    Specific Power2.04e-04 kg/MW

    Orion AKA "old Boom-boom" is the ultimate consumable nuclear thermal rocket, based on the "firecracker under a tin can" principle. Except the tin can is a spacecraft and the firecracker is a nuclear warhead.

    This concept has the spacecraft mounted with shock absorbers on an armored "pusher plate". A stream of small (5 to 15 kiloton) fission or fusion explosives are detonated under the plate to provide thrust. While you might find it difficult to believe that the spacecraft can survive this, you will admit that this will give lots of thrust to the spacecraft (or its fragments). On the plus side, a pusher plate that can protect the spacecraft from the near detonation of nuclear explosives will also provide dandy protection from any incoming weapons fire. On the minus side I can hear the environmentalists howling already. It will quite thoroughly devastate the lift-off site, and give all the crew bad backs and fallen arches. And they had better have extra-strength brassieres and athletic supporters.

    Mathematician Richard Courant viewed an Orion test and said "Zis is not nuts, zis is super-nuts."

    This section is about the Orion propulsion system. If you want all the hot and juicy details about various versions of Orion spacecraft go here.

    Please note that Orion drive is pretty close to being a torchship, and is not subject to the Every gram counts rule. It is probably the only torchship we have the technology to actually build today.

    If you want the real inside details of the original Orion design, run, do not walk, and get a copies the following issues of of Aerospace Projects Review: Volume 1, Number 4, Volume 1, Number 5, and Volume 2, Number 2. They have blueprints, tables, and lots of never before seen details.

    If you want your data raw, piled high and dry, here is a copy of report GA-5009 vol III "Nuclear Pulse Space Vehicle Study - Conceptual Vehicle Design" by General Atomics (1964). Lots of charts, lots of graphs, some very useful diagrams, almost worth skimming through it just to admire the diagrams.


    The dirty little secret about Orion is that the mission it is best suited for is boosting heavy payloads into orbit. Which is exactly the mission that the enviromentalists and the nuclear test ban treaty will prevent. Orion has excellent thrust, which is what you need for lift-off and landing. Unfortunately its exhaust velocity is pretty average, which is what you need for efficient orbit-to-orbit maneuvers.

    • LIFT-OFF: Orion is great! Can boost huge amounts of payload into orbit. Alas, nobody is going to allow hundreds of nuclear bombs to be detonated in Terra's atmosphere.
    • ORBIT-TO-ORBIT: Orion is run-of-the-mill average. There are other propulsion systems with much better exhaust velocities.
    • LANDING: DO NOT TRY! This will destroy the spacecraft, since you will be flying into the fireball of your nuclear detonations.

    Having said that, there is another situation where high thrust is desirable: a warship jinking to make itself harder to be hit by enemy weapons fire. It is also interesting to note that the Orion propulsion system works very well with the bomb-pumped laser weapons system.


    Each pulse unit is a tiny nuclear bomb, encased in a "radiation case" that has a hole in the top. A nuclear blasts is initially mostly x-rays. The radiation case is composed of a material that his opaque to x-rays (depleted uranium). The top hole thus "channels" the flood of x-rays in an upwards direction (at least in the few milliseconds before the bomb vaporizes the radiation case).

    The channeled x-rays then strike the "channel filler" (beryllium oxide). The channel filler transforms the atomic fury of x-rays into an atomic fury of heat.

    Lying on top of the channel filler is the disc of propellant (tungsten). The atomic fury of heat flashes the tungsten into a jet of ionized tungsten plasma, traveling at high velocity (in excess of 1.5 × 105 meters per second). This crashes into the pusher plate, accelerating the spacecraft. It crashes hard. You will note that there are two stages of shock absorbers between the pusher plate and the spacecraft, preventing instant crew death.

    The ratio of beryllium oxide to tungsten is 4:1.

    The thickness of the beryllium oxide and tungsten should be such to serve as a shield to protect the engine and upper vehicle from the neutron and high-energy gamma radiation produced by the nuclear explosion. This sets a lower limit on the thickness of the propellant and channel filler for a particular design.

    The jet is confined to a cone about 22.5 degrees (instead of in all directions). The detonation point is positioned such that the 22.5 cone exactly covers the diameter of the pusher plate. The idea is that the wider the area of the cone, the more spread out the impulse will be, and the larger the chance that the pusher plate will not be utterly destroyed by the impulse.

    It is estimated that 85% of the energy of the nuclear explosion can be directed in the desired direction. The pulse units are popped off at a rate of about one per second. A 5 kiloton charge is about 1,152 kg. The pulses are so brief that there is no appreciable "neutron activation", that is, the neutron from the detonations do not transmute parts of the spacecraft's structure into radioactive elements. This means astronauts can exit the spacecraft and do maintenance work shortly after the pulse units stop detonating.

    The device is basically a nuclear shaped charge. A pulse unit that was not a shaped charge would of course waste most of the energy of the explosion. Figure that 1% at best of the energy of a non-shaped-charge explosion would actually hit the pusher plate, what a waste of perfectly good plutonium.


    A short digression:

    When a nuclear device explodes, 90% of the bomb energy appears as electromagnetic radiation (80% soft X-rays and 10% gamma rays). So in airless space, a nuclear weapon destroys its target by x-rays.

    However, things are different inside Terra's atmosphere, or other planet's atmo. Atmospheres are typically opaque to x-rays (as Larry Niven put it: "so much for Superman's x-ray vision"). Which means the flux of x-rays are rapidly absobed, converting room-temperature air into a raging fireball with a temperature of roughly 100,000,000° Celsius. This is called blast. The end result is that nukes detonated inside an atmosphere are much more efficient at causing widespread destruction than nukes detonated in space.

    What does this mean? Orion pulse units designed to be used inside a planet's atmosphere can get away with using much smaller kiloton yield explosion sizes than units used in the vacuum of space. The atmosphere can also act as extra propellant.

    In the early Project Orion designs each charge was to accelerate the spacecraft by about 12 m/s. So in space a 4,000 ton spacecraft would use 5 kiloton charges, and a 10,000 ton spacecraft would use 15 kiloton charges. But during blast-off inside Terra's atmosphere, they could use 0.15 kt and 0.35 kt respectively. Quite a saving on plutonium.

    In the reports they only calculated the atmospheric effects for Terra's atmosphere. Because there are very few nearby planets and moons with atmospheres that are safe for space explorers to visit. You can see the different yield sizes in the Pulse Unit Table.


    How much weapons-grade plutonium will each charge require? As with most details about nuclear explosives, specifics are hard to come by. According to GA-5009 vol III , pulse units with 2.0×106 newtons to 4.0×107 newtons all require approximately 2 kilograms per pulse unit, with 1964 technology. It goes on to say that advances in the state of the art could reduce the required amount of plutonium by a factor 2 to 4, especially for lower thrust units. 2.0×106 n is 1 kiloton, I'm not sure what 4.0×107 n corresponds to, from the document I'd estimate it was about 15 kt. Presumably the 2 kg plutonium lower limit is due to problems with making a critical mass, you need a minimum amount to make it explode at all.


    According to Scott Lowther, the smallest pulse units were meant to propel a small ten-meter diameter Orion craft for the USAF and NASA. The units had a yield ranging from one-half to one kiloton. The USAF device was one kiloton, diameter 36 centimeters, mass of 86 kilograms, tungsten propellant mass of 34.3 kilograms, jet of tungsten plasma travels at 150,000 meters per second. One unit would deliver to the pusher plate a total impulse of 2,100,000 newton-seconds. Given the mass of the ten-meter Orion, detonating one pulse unit per second would give an acceleration well over one gee. According to my slide rule, this implies that the mass of the ten-meter Orion is a bit under 210 metric tons.

    Pulse UnitYieldMassDia.HeightPropellant
    per unit
    NASA 10m Orion
    141 kg0.86 s18,200 m/s
    (1,850 s)
    3.0×106 N3.5×106 N
    USAF 10m Orion
    1 kt79 kg
    (86 kg?)
    0.33 m0.61 m34.3 kg
    1 s1.5×105 m/s25,800 m/s
    (2,630 s)
    2.0×106 N2.0×106 N
    20m Orion
    450 kg0.87 s30,900 m/s
    (3,150 s)
    1.4×107 N1.6×107 N
    4000T Orion
    0.15 kt1,152 kg0.81 m0.86 m1.1 s1.17×105 m/s42,120 m/s
    (4,300 s)
    8.8×107 N8.0×107 N
    4000T Orion
    5 kt1,152 kg0.81 m0.86 m415 kg
    1.1 s1.17×105 m/s42,120 m/s
    (4,300 s)
    8.8×107 N8.0×107 N
    10,000T Orion
    0.35 kt118,000 m/s
    (12,000 s)
    10,000T Orion
    15 kt118,000 m/s
    (12,000 s)
    20,000T Orion
    29 kt1,150 kg0.8 m
    • Pulse Unit: The type of Orion spacecraft that uses this unit, and whether it is an atmospheric or vacuum type.
    • Yield: Nuclear explosive yield (kilotons)
    • Mass: Mass of the pulse unit
    • Dia.: Diameter of pulse unit
    • Height: Height of pulse unit
    • Propellant (percent): Mass of tungsten propellant in kilograms, as percentage of pulse unit mass in parenthesis.
    • Det. Interval: Time delay interval between pulse unit detonations.
    • Propellant Velocity: The velocity the tungsten propellant plasma travels at. Do not use this for delta V calculations.
    • Effective Exhaust Velocity (Isp): A value for exhaust velocity suitable for delta V calculations. Specific impulse in parenthesis.
    • Thrust per unit: Amount of thrust produced by detonating one pulse unit.
    • Effective Thrust: Thrust per second. Calculated by taking Thrust per unit and dividing by Det. Interval.

    Tungsten has an atomic number (Z) of 74. When the tungsten plate is vaporized, the resulting plasma jet has a relatively low velocity and diverges at a wide angle (22.5 degrees). Now, if you replace the tungsten with a material with a low Z, the plasma jet will instead have a high velocity at a narrow angle. The jet angle also grows narrower as the thickness of the plate is reduced. This makes it a poor propulsion system, but an effective weapon. Instead of a wall of gas hitting the pusher plate, it is more like a directed energy weapon. The military found this to be fascinating, who needs cannons when you can shoot spears of pure nuclear flame? The process was examined in a Strategic Defense Initiative project called "Casaba-Howitzer", which apparently is still classified. Which is not surprising but frustrating if one is trying to write a science fiction novel or spacecraft combat game.

    NASA has been quietly re-examining ORION, under the new name of "External Pulsed Plasma Propulsion". As George Dyson observed, the new name removes most references to "Nuclear", and all references to "Bombs."

    For details about spacecraft using Orion propulsion, go here.

    Oh, and another thing. ORION is fantastic for boosting unreasonably huge payloads into orbit and it is pretty great for orbit to orbit propulsion. But trying to use it to land is not a very good idea. At least not on a planet with an atmosphere.

    Project Orion
    Project Orion
    Exhaust Velocity19,620 m/s
    Specific Impulse2,000 s
    Thrust2,215,200 N
    Thrust Power21.7 GW
    Mass Flow113 kg/s
    Total Engine Mass203,680 kg
    Frozen Flow eff.39%
    Thermal eff.99%
    Total eff.39%
    Curium 245
    ReactorPulse Unit
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorPusher Plate
    Specific Power9 kg/MW

    This fabled technology converts the impulses of small nuclear detonations into thrust.

    The small shaped-charge bombs each have a mass of 230 kg (including propellant) and a yield of a quarter kiloton (1 terajoule). The fissile material is curium 245, with a critical mass of 4 kg, surrounded by a beryllium reflector. The soft X-rays, UV and plasma from the external detonation vaporize and compress the propellant to a gram per liter, highly opaque to the bomb energies at the temperatures attained (67000 K).

    The propellant, a mixture of water, nitrogen, and hydrogen, interfaces with a pusher plate “nozzle”, which can be either solid or magnetic.

    Shown is a solid plate, which tapers to the edges (to maintain a constant net velocity of the plate given a greater momentum transfer in the center). Pressure on the plate reaches 690 MPa in the center. The impulse shock is absorbed by a set of pneumatic “tires”, followed by gas-filled pistons detuned to the 56 Hz detonation frequency.

    The shock plate system becomes a useful shield if pointed towards the enemy.

    The amount of blast energy utilized for thrust is 7%, and the amount of pulse mass that intercepts the plate is 39%. A 56 TWth design optimized for 1TJ bombs achieves a specific impulse of 2 ksec and a thrust of 2.2 MN.

    Ted Taylor’s classic design, optimized for low yield bombs and 2 ksec specific impulse: “Project Orion”, George Dyson, Henry Holt and Company, 2002.

    From High Frontier by Philip Eklund

    Interesting e-mail conversation I had with Rhys Taylor on the topic of Entry-Descent-Landing (EDL) as relevant to nuclear pulse propulsion.

    I was aware one of the concepts that came out of the 1958 Project Orion involved landing a surface installation and a 100 man crew on the surface of Mars. Two of the early large Orion's would be involved. One would enter a low Mars orbit and completely cancel its orbital velocity while well above the sensible Martian atmosphere. The crew would ride down in a number of smaller landing craft with individual return stages. A large section of the vehicle, the base structure carrying a cargo of surface rovers, scientific gear, and consumables, would separate from the Orion propulsion module and descend propulsively on rockets without undergoing meteoric entry. The propulsion module would be allowed to crash on the surface (presumably this would entail transferring any remaining pulse units to the second Orion remaining in orbit before cancelling its orbital velocity — so only the absolute minimum required number of pulse units would remain to be expended before its uncontrolled descent and crash landing).

    My interest was in regards to soft landing an Orion intact after a controlled descent, and I was unsure of how deep into the atmosphere the nuclear pulse propulsion system could be fired, if it could be fired in descent mode, or if this was even advisable.

    Rhys was kind enough to advise me on these particular points, which to sum up are:

    1. Orion is capable of completely cancelling its orbital velocity.
    2. Descent would be a matter of managing the free-fall velocity of the vehicle.
    3. Inside the atmosphere the pulse unit will generate a many-thousands degree fireball, this is not a problem during launch, or in the vacuum of space, but during descent flying into the fireball would not be a good thing for vehicle and crew.
    4. There is some point at very high altitude where you would have to trade off from nuclear pulse propulsion to rocket powered descent.

    The input Rhys provided went toward this spacecraft designed for my Orion's Arm future history, and will be applied to several related spacecraft to be posted in the near future. 

    Orion Thrust and Isp

    For what it is worth, designs that I have seen in technical reports have specific impulses ranging from 3,354 seconds to 12,000 seconds (exhaust velocity 32,900 m/s to 12,000 m/s).

    Even though only a fraction of the pulse unit's mass is officially tungsten propellant, you have to count the entire mass of the pulse unit when figuring the mass ratio. The mass of the Orion spacecraft with a full load of pulse units is the wet mass, and the mass with zero pulse units is the dry mass.

    The thrust is not applied constantly, it is in the form of pulses separated by a fixed detonation interval. Generally the interval is from about half a second to 1.5 seconds. This means to figure the "effective" thrust you take the thrust-per-pulse-unit and divide it by the detonation interval in seconds. So if each pulse unit gives 2×106 Newtons, and they are detonated at 0.8 second intervals, the effective thrust is 2×106 / 0.8 = 2.5×106 Newtons

    Obviously the converse is if you have the effective thrust, you multiply it by the detonation interval to find the thrust-per-pulse-unit. So if the effective thrust is 3.5×106 N and the units are detonated at 0.86 second intervals, the thrust-per-pulse-unit is 3.5×106 N * 0.86 = 3.01×106 Newtons

    Bottom line: you can vary the Orion's thrust by varying the delay between detonations. The shorter the delay, the higher the effective thrust. Of course there is a lower limit on the delay, you have to wait for the first explosion to dissipate before initiating the second. Throwing a second pulse unit into an active nuclear explosion is just asking for a disaster.

    There are some interesting equations in GA-5009 vol III on pages 25 and 26 on the subject of nuclear pulse units. These were developed in the study for the 10 and 20 meter NASA Orion spacecraft, and they heavily rely upon a number of simplifying assumptions. These were for first generation pulse units, with the assumption that second generation units would have better performance. So take these with a grain of salt.

    These equations are only considered valid over the range 3×106 < FE < 2×108

    You are given the amount of thrust you want to get out of the propulsion system: FE and the detonation interval time Dp. From those you calculate the amount of thrust each pulse unit has to deliver Fp:

    Fp = FE / Dp

    From this the specific impulse, nuclear yield, and the mass of the Orion propulsion module.

    Isp = 1 / ((5.30×102 / (Fp * (1 + (2.83×10-3 * Fp1/3)))) + ((4.32×10-2 * (1 + (2.83×10-3 * Fp1/3))) / Fp1/3))

    Ve = Isp * g0

    Y = 9.30×10-10 * Fp4/3

    ME = Fp / (3.6 * g0)


    • FE = effective thrust (newtons)
    • Dp = delay between pulses (seconds)
    • Fp = thrust per pulse (newtons)
    • Isp = effective specific impulse (seconds)
    • Ve = exhaust velocity (m/s)
    • Y = size of nuclear yield in pulse unit (kilotons)
    • ME = mass of Orion propulsion module (kg)
    • g0 = acceleration due to gravity = 9.81 m/s2
    • x1/3 = cube root of x

    The results are close but do not exactly match the values given in the document, but they are better than nothing

    NASA 10-meter Orion
    Given Effective Thrust3.5×106 N
    Given Detonation Delay0.86 s
    Specific Impulse1,850 s1,830 s
    Yield1 kt0.4 kt
    Propulsion module mass90,946 kg85,245 kg
    NASA 20-meter Orion
    Given Effective Thrust1.6×107 N
    Given Detonation Delay0.87 s
    Specific Impulse3,150 s3,082 s
    Yield5 kt3.1 kt
    Propulsion module mass358,000 kg394,223 kg

    For more in depth calculations of an Orion rocket's specific impulse, read these two pages. But be prepared for some heavy math.



         There are three principal factors that control the effective specific impulse: (1) the mean propellant velocity, (2) the fraction of total propellant flow fc, which intercepts the pusher of the vehicle, and (3) a mass loss factor fm, which accounts for other mass necessarily ablated from the pusher plate of the vehicle, but which has been assumed herein to contribute little impulse to the vehicle in doing so.

         Based on the model used, the following conclusions were drawn:

    1. There is an optimum pulse energy for a given system (i. e., a given pusher diameter, and opacity of ablated pusher material) to yield a maximum specific impulse.
    2. Increasing the mean propellant velocity does not necessarily result in an increased effective specific impulse for a given system.
    3. Mean opacities of 103 square meters per kilogram or above appear necessary to approach the achievement of maximum effective specific impulses.
    4. Increasing the vehicle size (i. e., increasing pusher diameter) leads to higher values of fc but lower values of fm. The resulting effective specific impulse tends to increase if the pulse energy is kept at the optimum value.


         A schematic diagram showing the use of externally exploded pulse units to propel a space vehicle is shown in sketch (a). A pulse unit containing fusionable material plus propellant mass is ejected into position behind the vehicle and the pulse energy triggered. The propellant mass expands. A fraction of the propellant intercepts the pusher plate of the vehicle and transfers momentum and heat to the vehicle. The heat flux causes some ablation of the pusher surface. A succession of such pulses is continued until the desired total impulse for the mission is obtained.

         In the usual chemical or electric rocket case, all the propellant mass ejected from the vehicle contributes its momentum to the vehicle; thus, the effective specific impulse (the impulse per unit weight flow) is a function only of the mean propellant velocity:

    (All symbols are defined in the appendix. )

         As is apparent in sketch (a), not all the mass ejected from the vehicle in the nuclear pulse case is effective in contributing impulse to the vehicle. For the pulse case, the effective specific impulse is


    is the total impulse intercepted by the pusher. The weight of material (other than the propellant) which is lost from the vehicle per pulse and which is assumed to contribute no effective momentum to the vehicle is represented by Wa. In the analysis herein, the only such material considered is material ablated from the pusher plate. The mass loss factor is defined as

    The effective specific impulse for the pulse system can thus be expressed as


         To evaluate the specific impulse for this type of system, then, one has to look at the mean propellant velocity v that the system can tolerate, the effectiveness with which the mass of the pulse can be collimated so as to intercept the pusher plate of the vehicle, and the unavoidable mass losses in the system, particularly those lost through interaction of the high velocity propellant with the pusher surface.

         The analysis that follows considers the characteristics of the expanding propellant to derive conditions affecting the interaction between the propellant and pusher surface. The principal interactions affecting the performance limitation of the system are the rate of heat transfer leading to pusher ablation, stress limits on the pusher plate material, and pulse unit design as reflected in the maximum amount of collimation that is attainable.


         The propellant mass will be at a high temperature and pressure condition immediately following the pulse energy release. As this material expands into the vacuum of space, some recombinations of ionized and dissociated products will occur. Eventually, through expansion, the density drops to the point where interparticle collisions are relatively unimportant, and the material continues to expand in what is called a "self-similar" manner. The characteristic relation of this type of expansion for the case herein is

    In the treatment herein, it is assumed that all the energy absorbed by the propellant mass appears finally as kinetic energy with a Maxwellian distribution about the mean velocity v:

         Also, it is assumed that the condition for collisionless expansion is reached in a relatively short distance, compared to the dimensions of the system, so that the expansion products are effectively emitting from a point source.

         The density at any location and time is given by

    From equation (8), then, the other properties readily follow:

    Mass flow rate per unit area:


    Energy flow rate per unit area:

         If equation (6) is used, the various property relations obtained are as follows:

         Some plots of equations (12) to (15) are shown in figure 1 for a particular Ep/r3 value. The time scales are nondimensional in this figure, the reference time being tr = r/v. One can note, then, that the time at which the maxima in the various parameters occur (tm) is a direct function of the separation distance r. Also, the total time of pulse interaction is of the order of three times tm, so that pulse interaction time also varies directly with separation distance. This fact has a bearing on the heat-transfer effects as will be seen later.

         Consider the arrival of propellant from a pulse onto a plane normal to the main flow direction (sketch (b)). The mass flux at point (z, β) normal to the plane is

    and the pressure on the plane at that point is

    Typical radial variations of pressure and density are shown in figure 2. The impulse per unit area on the plane at (z, β) in the z-direction is

    The total impulse in the z-direction from flux within a given cone angle θ is

    But the fraction of the total mass flow which is included within a given cone angle θ is

    for isotropic distribution. Thus, considering only the propellant interception factor, the effective specific impulse of a system using a circular pusher plate to intercept the mass flow within a cone angle θ is

         If a pulse unit is designed to direct a disproportionate fraction of the total mass into a given cone angle, the flow is still assumed uniformly distributed within the cone. The parameter C (hereafter referred to as a collimation factor) is defined as the ratio of the enhanced total mass in the cone to the amount in the cone if the distribution were isotropic. It can be expressed as

    With this assumption of the flow distribution, the relations developed for the isotropic distribution may be used for collimated flow cases by replacing the propellant mass term Mp with the product C Mp .

         If the case with collimated flow is now assumed, the relation for total impulse (eq. (19)) is put in terms of pulse energy and C:

    The pressure is highest at the center of the plate (β = 0):

    At β = 0, the maximum pressure at any time is

         The effect of two factors (equivalent energy of the pulse and attainable collimation factor) involved in the design of pulse units on the attainable propellant interception factor fc can be seen by the following development. Combining equations (21)and (23) yields

    In all subsequent relations where a specific value of maximum pressure is used, a value of pm = 6.9×108 newtons per square meter (equivalent to about 100 000 psi) is used. This value represents a reasonable upper limit to the allowable yield stress of materials that might be used for the pusher. If this value is used, equation (25) becomes

         The pusher diameter required to intercept the flux in a given cone angle θ is obtained from the geometric relation (shown in sketch (c)).

    Combining equations (26) to (28) yields the relation

    With the assumption that all the propellant mass has the mean velocity v, the upper limit to fc is 0.5 in order to satisfy the momentum balance requirement. Equation (29), with the upper limit restriction of 0.5 for fc, is shown plotted in figure 3 for four values of total pulse energy, Ep = 4.18×109, 4.18×1010, 4.18×1011, 4.18×1012 joules (1, 10, 100, and 1000 ton equivalents). (For reference, 1 gram of deuterium-tritium (DT) mixture fully reacted is equivalent to about 3.344×1011joules (80 tons TNT). ) The interrelation among vehicle size (as reflected by pusher diameter), pulse energy size (Ep) and pulse unit design (as reflected by the collimation factor C) is evident.

    When the separation distance is maintained at the smallest value permitted by pressure limitations (the conditions imposed for fig. 3), one can note that there is a large improvement in the propellant interception factor fc for (1) better collimation of the propellant, (2) smaller energy pulse value for the same degree of collimation, and (3) larger size vehicles (i. e., larger pusher diameters).


         When the flux of propellant arrives at the pusher plate, the initial, high velocity particles cause some sputtering. Some penetration into the material of the pusher also occurs. The effect of this initial bombardment is small compared to the eventual ablation caused by the arrival of the remainder of the pulse mass. The propellant that arrives is assumed to just lose its kinetic energy and form a hydrodynamic stagnation layer. The temperature of this stagnation layer is calculated by assuming that it reaches equilibrium through blackbody radiation back to the vacuum of space.

         Thus, when the energy flow rate relation (15) is used, the equilibrium can be expressed as


         A plot of this relation is shown in figure 4. Stagnation temperatures up to the 20- electron-volt range are typical.

         The formation of this high temperature layer causes the temperature of a pusher surface to rise quickly to the ablation level. The ablated gas then forms a protective layer and slows down subsequent ablation rates.

         At the temperature levels of the stagnation layer of gas, heat transfer to the pusher is mainly by radiation. Therefore, a surface material whose ablated products have a high absorptance for radiation in the frequency range characteristic of the temperature of the stagnation layer would be a distinct advantage in limiting ablation amounts. This situation is similar to that encountered in the gas-core nuclear rocket.

         In order to estimate the magnitude of the heat-transfer problem, the model of the interaction process shown in sketch (d) was assumed. The ablated vapor layer and the high-temperature layer from stagnating propellant were assumed to remain unmixed. The flow of heat through the ablated vapor layer was then looked at as a combined process of radiation and conduction. It is assumed that the ablated vapor, of necessity, will be of material which has a high absorptance for radiant energy at the temperature of the stagnation layer. The one-dimensional equation describing the transfer of heat in an optically thick radiative-conductive medium with no heat source terms is

    where aR is the adsorptance of vapor layer layer.

         Numerical solutions to equation (32) may be obtained by assuming a boundary temperature history at x = 0 (given by the stagnation temperature relation (31)) and with the temperature gradient approaching zero as x becomes large. Typical temperature profiles for these solutions are shown in figure 5. The rate of advance of the temperature "front" into the ablated vapor medium is also obtained.

         The inflection point of the temperature-distance relation occurs at about the temperature level for which equal rates of heat transfer occur by both radiation and conduction modes, This condition is given approximately by

    This relation becomes

    when substituting for σ and expressing temperature in electron volt units. The temperature level of the major portion of the ablated medium is below this inflection point temperature. This observation is noted here because of its relation to the properties of the ablated layer.

         The assumption that most of the radiant energy from the stagnation layer is absorbed in a relatively small thickness of the ablated vapor leads to the temperature profiles noted in figure 5, that is, a relatively steep temperature front that propagates through the medium. The similarity here to the gas-core rocket profiles is noted. In the gas-core case, seeded gas is fed toward the advancing temperature front at such a rate as to maintain the steep temperature profile away from the wall of the rocket chamber. In the pusher case herein, ablated material is fed into the vapor state at a rate governed by the amount of heat that reaches the wall and the amount of energy required to vaporize plate material and raise it to the temperature level of the ablated medium.

         In the following calculation of the amount of pusher ablation, it is assumed that heat transfer to the surface by radiation (though only a small fraction of the total radiant energy available) is still the major mode of energy transfer. The incremental heat transfer per unit area, then, is

    The increment of mass ablated per unit area is

         Equations (35) and (36) were solved numerically for energy input using equation (30) when no ablated layer (x = 0) was present at t = 0 and when any decrease in x(t) resulting from propagation of the temperature front into the ablated medium was neglected. The calculated ablation amounts will thus be lower than actual because of this latter assumption.

    Typical total ablation per unit area values are shown in figure 6. These were calculated using a value of 5×107 joules per kilogram for Ha.

    The total ablation amount, however, is not especially sensitive to Ha as shown in figure 7. A factor of 10 change in Ha causes only a 15 to 20 percent change in the total ablation amount under the conditions of these calculations.

    Figure 8 shows the range of energy absorption that goes into just ionization and thermal energy of the species for two metal vapors, iron (Fe) and uranium (U). Sublimation energy is, of course, also included in Ha.

         The effect of separation distance on ablation rates is shown infigure 9 by a comparison of curves. As the separation distance is increased, the energy arrival rate (per unit area) decreases. However, the total interaction time increases. The net effect (fig. 9) of increasing the separation distance to four times the minimum separation distance (eq. (28)) was to decrease the total ablation by only about 32 percent at Ep = 4.18×1012 joules (1000 tons) and 38 percent at Ep = 4.18×109 joules (1 ton).

         The ablation calculations shown up to this point have been for conditions at the center of the pusher (β = 0). Away from the center (β > 0), the ablation decreases as shown in figure 10. The same countering influence of the two factors, energy intensity and interaction time, are present in the radial variation as in the separation distance variation (fig. 9). For Ep = 4.18×109 joules (1 ton), a 17 percent decrease in the ablation rate was calculated at 45° off the axis.

         Of special interest now is a calculation of the mass loss factor fm. Since the only such loss considered herein is that by means of ablation from the pusher surface, fm is given by equation (4).

         A cylindrical pusher plate of diameter d (sketch (e)) is now considered. The ratio of total ablation per pulse to propellant mass for the pusher plate of diameter d is

    The energy arrival rate at (z, β) for a pulse with collimation factor C is

    Now, let tz = z/v . Then tr = tz sec β and, using equation (16),

    At this point the calculations are restricted to values of z corresponding to the maximum pressure limitation (eq. (27)). When equation (27) is used, equation (39) becomes

    Equation (37) is transformed to

    by using the previous relations and putting the pulse energy in terms of Ep.

         The inner integral (with respect to t/tz) was solved numerically under the same assumptions as used in solving equations (35) and (36). The resulting values of the mass loss factor fm for several combinatims of variables are shown in figure 11. The curves stop at the limiting diameter where the propellant interception factor is 0.5.

         The larger the pusher diameter, the smaller is the factor fm, other factors being constant. More pusher area is exposed to ablating conditions, while pulse mass Mp remains constant. The mass ablation loss factor is lower for smaller energy pulses for the same size pusher plate. This is because the separation distance zm is lower for lower energy pulses, and the pusher intercepts a greater fraction of the total energy flux from the pulse.

         The absorption coefficient aR has a marked effect on the mass loss factor since, as was noted before (fig. 6), the total ablation amount is nearly inversely proportional to aR for the conditions of these calculations.


         A simplified model of the overall processes involved in the external nuclear pulse propulsion scheme has been assumed. Only the features of the propulsion system that affect the overall or effective specific impulse have been considered. The propellant flow from the pulse energy source has been assumed to occur self-similarly from a point source. A collimation factor has been used to account for the concentration of more propellant into smaller flow cone angles than would occur simply from isotropic expansion. The hydrodynamic stagnation layer temperature is determined by the balance between incoming kinetic energy and blackbody radiation back to space. Ablation of pusher surface material occurs by radiant heat transfer from the stagnation temperature through the ablated layer. A mean opacity for the ablated layer material is assumed.

         Using this model, two principal factors determining the effective specific impulse were calculated: the propellant loss factor fc (eq. (29) and fig. 3), and the mass loss factor fm (eq. (4) and fig. 11). These two factors are combined to yield the ratio

    Some typical plots of relation (42) are shown in figure 12 for a collimation factor of 3 and aR value of 102 square meters per kilogram. It is apparent in this figure that the optimum specific impulse for smaller pusher diameters is attained with the small pulse energies. Under the conditions of figure 12 there is little or no improvement in effective specific impulse in going to higher mean propellant velocities at the smaller pulse energy levels.


          Dear Mr. Taylor
         I am making a Project Orion mod for Kerbal Space Program.
         Some people have expressed a desire to vary the thrust of the propulsion system. From my reading of the 1960's era documents, they proposed doing this by altering the frequency of the detonations.
         This means the impulse imparted to the spacecraft per detonation is constant, but by increasing the time, the effective acceleration is reduced.
         But the ship will still get kicked by the full impulse.
         Somebody mentioned that in Larry Niven and Jerry Pournelle's novel FOOTFALL, the orion drive vessel would vary the impulse per bomb by varying the standoff distance between the bomb detonation point and the bottom of the pusher plate.
         If you have the time, I'd like your thoughts on the feasibility of that idea.
         Offhand I'd say that by increasing the standoff distance, you decrease the amount of the propellant and/or bomb blast that actually strikes the pusher plate. This would lower the impulse per bomb.
         Decreasing the standoff distance would have no effect on the amount of blast striking the pusher plate, it would all hit the plate, but just in a more concentrated area of the plate. Which would probably severely damage the plate.
         Footfall also mentions using a form of thrust vectoring by detonating the bombs off-center. This strikes me as a very bad idea. The impulse would then not be delivered co-axially with the shock absorber shafts, probably doing severe damage.

    — Winchell Chung

         From what I remember, the yield of the bombs would vary during launch from 0.1 kt at low altitude to 5-15 kt above the atmosphere I don't recall if that actually alters the thrust or not, because the extra mass swept up by the air will make a big difference. But certainly that would be one way to do it in a vacuum.

         For a carefully controlled launch, that's one thing. You'd just pre-arrange the bombs so that they're ejected in the right order. As for selecting bombs of the appropriate yield on the fly, I don't know. It seems like adding a lot of complicated moving equipment which will have to work perfectly otherwise all is lost (unless you have a Star-Trek like ability to control the yield of individual bombs electronically, on demand). But maybe that isn't so much extra crazy compared to the rest of the project.

         Altering the detonation frequency will work during a launch phase — longer intervals will mean more deceleration due to gravity. I can't see that makes any difference in an orbital environment though.

         Increasing the standoff distance seems like the best approach to me. In an orbital environment I can't immediately see any problems with this. But I don't think it works as a catch-all solution : above the atmosphere, you absolutely need the higher yield bombs to deliver a useful specific impulse. At low altitudes, you really aren't going to want to detonate a 5kt nuke anywhere nearby. A more powerful gas gun could shoot them to greater distances to lower the damage, but we're still talking about something that's effectively a WMD pointed the wrong way.

         I agree that decreasing the standoff distance doesn't help beyond the point where the pusher plate is able to absorb the majority of the plasma. The caveat is how directional the nukes are, which of course we can't know precisely. There was the Medusa concept of using a sail to catch more of the plasma, and I think the ship in Footfall had a curved plate. But it sounded to me that getting quite a tight beam wasn't that much of a problem.

         Perhaps one solution to allow variable thrust in all environments is to have two different avaiable yields for the pulse units. In vaccum, you have to use the high yield devices. For atmosphere, choose a lower yield but something that's higher than the nominal 0.1 kt that you could get away with in the denset part of the atmosphere. Compensate by varying the standoff distance and detonation interval, so that you avoid delivering the maximum impulse at ground level (which would be too high to withstand). Gradually lower the distance and reduce the interval until you're delivering the maximum impulse, then switch to the higher yield devices (if necessary, starting those with a higher than optimum standoff distance and detonation interval too). I don't have any numbers to say if this is plausible or not though.

         I also agree that off-center detonations reek of suicide. Given the massive accelerations (something like 1000g on the pusher plate) I can't see how this wouldn't end badly (I don't know for certain though, I never even tried the math on this one). I prefer the Mythbusters solution of "strap rockets onto everything" and rotating the ship in between detonations.

    (ed note: Aldo Spadoni had some thoughts about the Footfall's ship's off-center detonations)

    by Rhys Taylor (2013)

    Orion Environmental Impact

    Naturally, some people freak out when you tell them about a rocket that rises into orbit by detonating Two! Hundred! Atom! Bombs!. But it actually isn't quite as bad as it sounds.

    First off, these are teeny-tiny atom bombs, honest. The nuclear pulse units used in space will be about one kiloton each, while the Nagasaki device was more like 20 kt. And in any event, the nuclear pulse units used in the atmosphere are only 0.15 kt ( about 1/130th the size of the Nagasaki device). This is because the atmosphere converts the explosion x-rays into "blast", increasing the effectiveness of the pulse unit so you can lower the kilotonnage.

    So we are not talking about zillions of 25 megaton city-killer nukes scorching the planet and causing nuclear winter.

    Some environmentalists howl that Orion should never be used for surface-to-orbit boosts, due to the danger of DUNT-dunt-Dunnnnnnnn Deadly Radioactive Fallout. However, there is a recent report that suggests ways of minimizing the fallout from an ORION doing a ground lift-off (or a, wait for it, "blast-off" {rimshot}). Apparently if the launch pad is a large piece of armor plate with a coating of graphite there is little or no fallout.

    By which they mean, little or no ground dirt irradiated by neutrons and transformed into deadly fallout and spread the the four winds.

    There is another problem, though, ironically because the pulse units use small low-yield nuclear devices.

    Large devices can be made very efficient, pretty much 100% of the uranium or plutonium is consumed in the nuclear reaction. It is much more difficult with low-yield devices, especially sub-kiloton devices. Some of the plutonium is not consumed, it is merely vaporized and sprayed into the atmosphere. Fallout, in other words. You will need to develop low-yield devices with 100% plutonium burn-up, or use fusion devices (with 100% burn-up fission triggers or with laser inertial confinement fusion triggers).

    The alternative is boosting the Orion about 90 kilometers up using a non-fallout chemical rocket. Which more or less defeats the purpose of using an Orion engine in the first place. Remember that Orions are best at boosting massive payloads into orbit.

    Most of the fallout will fall within 80 kilometers of the launch site. You can also reduce the fallout by a factor of 10 if you launch from near one of the two Magnetic Poles. You see, far from the magnetic poles, Terra's magnetic field traps fallout particles that would have been ejected into space, and returns them to the surface. At the magnetic poles are "holes" in the magnetic field which allows the fallout to travel unimpeded into deep space.

    Annoyingly this is real hard to do because the freaking magnetic pole keeps wandering around.

    One minor drawback is that if you launch from a magnetic pole, you pretty much have to launch into a polar orbit. In practice these are seldom used specialized orbits, of use mainly for military spy satellites, weather satellites, orbital bombardment weapons, and Google Earth. The Orion will probably have to change to a more useful equatorial orbit, which alas will require a change-of-plane maneuver of ninety freaking degrees. COPMs are notorious for being the most costly all maneuvers in terms of delta-V, and that is for changes of only a few degrees. This is still only a minor drawback because an Orion has delta-V to burn. It can do a 90° COPM and not even notice the delta-V is missing. As Jeff Zugale says: "Pretty sure that’s a rounding error when using nukes to launch 5000 tons." So this is a case of Crazy nastya$$ Orions just don't give a sh*t.

    When fissionables like plutonium undergo fission, their atoms are split which produces atomic energy. The split atoms are called fission fragments.

    The good news is that they have very short half-lives, e.g., in 50 days pretty much all of the Strontium 94 has decayed away (because 50 days is 58,000 St94 half-lives).

    The bad news is that they have very short half-lives, this means they are hideously radioactive. Radioactive elements decay by emitting radiation, shorter half-life means more decays per second means a higher dose of radiation per second.

    The fragments that come screaming out of the detonation aimed at the sky are no problem. They are moving several times faster than Terra's escape velocity, you will never see them again (Terra's escape velocity is 11.2 km/s, the fragments are travelling like a bat out of hell at 2,000 km/s). The ones aimed towards Terra are a problem. The fragments can be reduced by using fusion instead of fission pulse units. The fragments can also be reduced by designing the pulse units to trade thrust in favor of directing more of the fragments skyward.

    A more sophisticated objection to using Orion inside an atmosphere is the sci-fi horror of EMP melting all our computers, making our smart phones explode, and otherwise ruining anything using electricity. But that actually is not much of a problem. EMP is not a concern unless the detonation is larger than one megaton or so, Orion propulsion charges are only a few kilotons (one one-thousandth of a megaton). Ben Pearson did an analysis and concluded that Orion charges would only have EMP effects within a radius of 276 kilometers (the International Space Station has an orbital height of about 370 kilometers). So just be sure your launch site is in a remote location, which you probably would have done anyway.

    Naturally watching an Orion blast-off is very bad for your eyes, defined as instant permanent blindness. This is called "eyeburn". While the Orion is below 30 km you definitely need protective goggles or you might be blinded. Above 90 km your eyesight it safe. In between 30 and 90 is the gray area, where prudent people keep their protective goggles on.

    Detonating pulse units in space near Terra can create nasty artificial radiation belts. The explosion can pump electrons into the magnetosphere, creating the belt.

    There are two factors: detonation altitude from Terra's surface, and magnetic latitude in Terra's magnetic field. If the detonation is within 6,700 kilometers of Terra's surface (i.e., closer than 2 Terran radii from Terra's center) and at a magnetic latitude from 0° to 40°, the radiation belt can last for years. Above 2 Terran radii the radiation belt will last for only weeks, and from latitude 80° to 90°, the radiation belt will last for only a few minutes.

    The military discovered this the hard way with the Starfish Prime nuclear test. The instant auroras were very pretty. The instant EMP was very scary, larger than expected (but the test was using a 1.4 megaton nuke, not a 0.001 megaton pulse unit). The artificial radiation belt that showed up a few days later was a very rude surprise. About one-third of all low orbiting satellites were eventually destroyed by the radiation belt.

    The radiation belts are harmless to people on Terra, but astronauts in orbit and satellites are at risk.

    There are three classes of pulse unit failure modes. Note that in this analysis the USAF had given up and had decided to boost the Orion on top of a chemical rocket.

    Class I - Pad Abort
    Typically occurs when the chemical booster burns or explodes on the pad. There will be no nuclear explosion. The pulse units contain chemical explosives, but there is much more explosive potential in the chemical booster fuel. Even if all the pulse units exploded simultaneously there would only be a 1 psi overpressure out to 300 meters and shrapnel hazard out to 2,000 meters.

    A chemical booster burn could aerosolize radioactive plutonium from booster units and create a downrange fallout hazard. The solution is to put the launch pad over a pool of water about 10 meters deep. In event of fire, collapse the pad into the pool. The fire would be extinguished and any escaped plutonium will be contained in the water. Many of the pulse units can be recovered and reused.
    Class II - Failure to Orbit
    The trouble is that the thousands of nuclear pulse units will fall down, probably into uncontrolled territory. As with Class I there will be no nuclear explosion, the chemical explosion will be impressive but not too huge, and there is a danger of radioactive fallout. All in what could very well be a foreign country.

    In addition, it will be scattering thousands of containers of weapons grade plutonium in convenient form to cause nuclear weapon proliferation. Or the pulse units could be used as is as impromptu terrorist devices. Though I'm sure the devices will contain fail-safes seven ways to Sunday, the same way nuclear warheads are in order to deal with the possibility of them falling into the Wrong Hands.

    Probably the best solution is to command all of the nuclear charges to detonate simultaneously while the spacecraft is at high altitude. This will make one heck of a fireworks display, and may cause an EMP, but nuclear devices in questionable hands is to be avoided at all costs.
    Class III - Misfire
    If a given pulse unit fails to detonate, the command can be resent repeatably, and/or there can be an automatic on-board destruct system. Otherwise the unit could survive reentry (due to the tungsten propellant plate) causing some damage to the country it hit and causing a foreign policy nightmare to the nation owning the Orion spacecraft.

    By about 1963 General Atomic had given up on designing an Orion to lift off from Terra's surface under nuclear power. They put together three plans for using chemical rocket boosters to get the Orion into orbit. Again this is throwing away the big advantage of the Orion, its ability to boost massive payloads.

    Mode I
    A fully loaded and fully fueled Orion is boosted to an altitude of 90 kilometers and 900 m/s by a chemical rocket. There it stages, and the Orion proceeds into orbit or into mission trajectory under nuclear power. The disadvantage is it requires a subobital start-up of the Orion engine. The Orion engine will need a thrust greater than the mass of the spacecraft, the standard was T/W of 1.25. But high thrust is never a problem with Orion.
    Mode II
    An empty Orion is loaded with just enough pulse units. It is boosted to an altitude of 90 kilometers and 900 m/s by a chemical rocket. There it stages, and the Orion proceeds into orbit. A second chemical booster rendezvous with the Orion to deliver the payload and a full load of pulse units.
    This was the worst plan. It combines the disadvantage of Mode I (by requiring suborbital start-up of the Orion engine) with the disadvantage of Mode III (by requiring orbital assembly).
    Mode III
    The Orion is boosted into orbit piecemeal as payload on a series of chemical boosters. The Orion is assembled in orbit, then departs on its mission under nuclear power. The main advantage is it avoids the possibility of the entire Orion spacecraft crashing to Terra in the event of a propulsion failure. The second advantage is it allowed a lower thrust Orion unit to be used, but with Orion thrust is never a problem. The main disadvantage is that orbital assembly is time consuming and difficult.


    Zeta pinch is a type of plasma confinement system that uses an electrical current in the plasma to generate a magnetic field that compresses it. The compression is due to the Lorentz force.

    Zeta-Pinch Fission

    Mini-Mag Orion
    Mini-Mag Orion
    Exhaust Velocity157,000 m/s
    Specific Impulse16,004 s
    Thrust1,870,000 N
    Thrust Power0.1 TW
    Mass Flow12 kg/s
    Total Engine Mass199,600 kg
    Curium 245
    Specific Power1 kg/MW
    Curium 245
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Mini-Mag Orion (DRM-1)
    Exhaust Velocity93,164 m/s
    Specific Impulse9,497 s
    Thrust642,000 N
    Thrust Power29.9 GW
    Mass Flow7 kg/s
    Total Engine Mass119,046 kg
    Wet Mass731,924 kg
    Dry Mass250,300 kg
    Mass Ratio2.92 m/s
    ΔV99,967 m/s
    Specific Power4 kg/MW
    Mini-Mag Orion (DRM-3)
    Exhaust Velocity93,000 m/s
    Specific Impulse9,480 s
    Thrust642,000 N
    Thrust Power29.9 GW
    Mass Flow7 kg/s
    Total Engine Mass199,600 kg
    Wet Mass788,686 kg
    Dry Mass157,723 kg
    Mass Ratio5.00 m/s
    ΔV149,686 m/s
    Specific Power7 kg/MW

    The Mini-MagOrion is a sort of micro-fission Orion propulsion system. The idea was to make an Orion with weaker (and more reasonably sized) explosive pulses, using pulse charges that were not self contained (the full Orion pulse units were nothing less than nuclear bombs). Subcritical hollow spheres of curium-245 are compressed by a Z-pinch magnetic field until they explode. The sacrificial Z-pinch coil in each pulse charge is energized by an a huge external capacitor bank mounted in the spacecraft. So the pulse units are not bombs.

    The explosion is caught by a superconducting magnetic nozzle.

    More details are in the Realistic Designs section.

    Z-pinch Microfission
    Z-pinch Microfission
    Z-pinch Microfission
    Exhaust Velocity156,960 m/s
    Specific Impulse16,000 s
    Thrust8,500 N
    Thrust Power0.7 GW
    Mass Flow0.05 kg/s
    Total Engine Mass193,333 kg
    Frozen Flow eff.74%
    Thermal eff.90%
    Total eff.67%
    Curium 245
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Specific Power290 kg/MW

    Electrodynamic zeta-pinch compression can be used to generate critical mass atomic bombs at very low yields. These detonations can be used to generate impulsive power or thrust.

    Exotic fission material (245Cm) is utilized to reduce the required compression ratio. The explosion of each low yield (335 GJ) atomic bomb energizes and vaporizes a set of low mass transmission lines, used to pump either another high current Z-pinch, or a bank of nanotube-enhanced ultracapacitors.

    Each bomb uses 40 grams of Cm fissile material and 60 grams of Be reflector material, with an aspect ratio of 5. A DT diode is used as a neutron emitter. The mylar transmission lines have a mass of 15 kg, and are replaced after each shot.

    The design illustrated is rated for a shot every 5.5 minutes, equivalent an output of 1000 MWth. If utilized for thrust, this provides 7.7 kN at a specific impulse of 17 ksec.

    Ralph Ewig & Dana Andrews, “Mini-MagOrion Micro Fission Powered Orion Rocket”, Andrews Space & Technology, 2002.

    From High Frontier by Philip Eklund
    n-Li6 Microfission
    n-Li6 Microfission
    n-6Li Microfission
    Exhaust Velocity156,960 m/s
    Specific Impulse16,000 s
    Thrust20,000 N
    Thrust Power1.6 GW
    Mass Flow0.13 kg/s
    Total Engine Mass106,667 kg
    Frozen Flow eff.87%
    Thermal eff.90%
    Total eff.78%
    Lithium 6
    ReactorUltracold Neutron
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Specific Power68 kg/MW

    The minimum explosive yield for fission bombs is about a quarter kiloton. Thus, rockets that fly using atomic explosions, such as Project Orion, require huge shock absorbers.

    The pulse energy can be brought down to microfission levels by the use of exotic particles. A n-6Li microfission thruster brings the lithium isotope 6Li to spontaneous microfission by interaction with particles with very large reaction cross sections such as ultracold neutrons. No “critical mass” is required. This clean reaction produces only charged particles (T and He), each at about 2 MeV.

    The system illustrated uses a 5-meter magnetic nozzle to transfer the microexplosion energy to the vehicle. This magnetic impulse transfer is borrowed from the MagOrion concept (combination of Orion and the magnetic sail).

    A fuel reaction rate of 60 mg/sec yields 3720 MWth. At a pulse repetition rate of one 224 GJ (0.05 kT) detonation each minute, the thrust is 12.8 kN at a 12 ksec specific impulse. A hydraulic fixture oscillates at a tuned frequency to provide a constant acceleration to the spacecraft. The combined frozen-flow and nozzle efficiencies are 21%, and the thermal efficiency is 96%.

    Ralph Ewig’s “Mini-magOrion” concept, modified for n-6Li fission,

    From High Frontier by Philip Eklund
    Ultracold Neutrons

    Neutrons are normally unstable particles, with a half life of 12 minutes.

    When polarized and ultra-cooled (using vibrators or turbines), they form a dineutron or tetraneutron phase. These “molecules” are believed to be stable and storable in total internal reflection bottles, lined with diamond-like carbon as the neutron reflector.

    Ultracold neutrons (UCN) have a huge quantum mechanical wavelength as a consequence of their slow movement (typically 0.4 μm @ 1 m/sec), and thus can spontaneously initiate fission reactions such as n-235U or n-6Li.

    If the neutron source is a nuclear reactor, the neutrons must be cooled from 2 MeV to 2 meV using a heavy water moderator, and then in a UCN turbine to 0.2 IeV.

    Robert L. Forward, “Alternate Propulsion Energy Sources”, 1983.

    From High Frontier by Philip Eklund

    Zeta-Pinch Fusion

    HOPE Z-Pinch
    Propulsion SystemZ-Pinch Fusion
    Exhaust Velocity189,780 m/s
    Specific Impulse19,346 s
    Thrust38,120 N
    Thrust Power3.6 GW
    Mass Flow0.20 kg/s
    Total Engine Mass95,138 kg
    FuelDeuterium-Tritium fusion
    + Lithium6 fission
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Wet Mass888,720 kg
    Dry Mass552,000 kg
    Mass Ratio1.61 m/s
    ΔV90,380 m/s
    Specific Power26 kg/MW
    Firefly Starship
    2013 design
    ΔV2.698×107 m/s
    Wet Mass17,800 metric tons
    Dry Mass2,365 metric tons
    Mass Ratio7.526
    Payload150 metric tons
    PropulsionZ-Pinch DD Fusion
    Exhaust Velocity1.289×107 m/s
    Thrust1.9×106 N
    Acceleration0.11 m/s
    (0.01 g)
    Accel time4 years
    Coast time93 years
    Decel time1 years

    PuFF Pulsed Fission Fusion

    Pulsed Fission-Fusion
    20,000 sec
    196,000 m/s
    Thrust29,400 N
    2.88 GW
    96 kW/kg
    U-235 + D-T

    This is from The Pulsed Fission-Fusion (PuFF) Propulsion System and Phase I Final Report.

    The study authors were going to take a Hope Z-Pinch Fusion spacecraft and swap out its drive for the PuFF drive.

    The idea is that while you can make some fuel undergo nuclear fission, and you can make other fuel undergo nuclear fusion, wouldn't it be nice to make some fuel do both? After all, a standard nuclear fusion warhead is a slug of fusion fuel that is ignited by the detonation of a small nuclear fission warhead.

    Refer to the diagram at right.

    The target is a charge of fission/fusion fuel, composed of Uranium-235 fission fuel and Deuterium-Tritium fusion fuel. The charge is held at the ignition point by some strong holder.

    A ring of liquid lithium sprayers (Li Injectors) are aimed at the target. They spray a cone-shaped plume of liquid lithium (Li Shell) with the cone apex located at the target. Oh, did I mention that the sprayers are connected to the anode of the power system capacitor (LTDs) so they and the lithium shell are charged to two mega-volts? The target holder is connected to the cathode.

    When the liquid lithium hits the target the circuit is closed, and the target is electrocuted by two mega-amps at two mega-volts (also totally draining the power system capacitor). This is 4 terawatts (4×1012 watts). Lorentz force (j×B) produced by the current and magnetic field savagely squeezes the fuel charge to one-tenth its original size. This makes the uranium achieve criticality.

    Only some of the uranium undergoes nuclear fission like an atom bomb (which it is). This heats the D-T fuel hot enough to initiate nuclear fusion.

    Neutrons from the fusion reaction ignites more of the uranium into a fission reaction. The heat from the fission boosts the fusion rate. Rinse-Lather-Repeat. This is called a Fission-Fusion Cascade. The fission to fusion cycle keeps cascading until all the fuel is burnt.

    The energy from the cascade turns the liquid lithium into plasma. The plume of charged plasma from the cascade is ejected by the magnetic exhaust nozzle. In addition to creating thrust, the nozzle also harvests some of the exhaust energy to charge up the primary power system capacitors for the subsequent pulse.

    Each fuel charge detonation takes several microseconds to cascade to full burnout. Detonations are repeated up to a rate of 100 Hz. The report notes that much analysis and experimentation is needed to find the optimum detonation frequency and fuel charge size.

    The specific impulse and thrust can shift gears by modifying the amount of lithium injected.

    Initially the power system capacitors are empty. For the first charge of the new burn an onboard SP-100 nuclear reactor laboriously charges them up. Subsequent capacitor recharges are by harvesting exhaust energy.

    Left as an exercise for the reader is what the heck do you make the target holder out of so it is not obliterated by the fission and fusion explosions.

    A - Target
    Charge of fission/fusion fuel
    B - Linear Transformer Drivers (LTD)
    Pulsed power storage (capacitors), discharge, and compression system
    C - Magnetic Nozzle (MN)
    Directs fission/fusion products into exhaust for thrust. Recovers energy for next pulse.
    D - Recharge System
    Pulse generation and onboard power storage/generation
    E - Lithium Injectors
    Lithium tankage / distribution system to provide target liner (cone of liquid lithium) and power conduction path (when it touches the target)
    F - Target Storage / Dispenser
    Maintains targets in non-critical configuration (so the uranium doesn't explode prematurely), injects into nozzle


    Exhaust velocity490,000 m/s
    to 980,000 m/s

    Medusa is driven by the detonation of nuclear charges like Orion, except the charges are set off in front of the spacecraft instead of behind. The spacecraft trails behind a monstrously huge parachute shaped sail (about 500 meters). The sail intercepts the energy from the explosion. Medusa performs better than the classical Orion design because its pusher plate intercepts more of the bomb's blast, its shock-absorber stroke is much longer, and all its major structures are in tension and hence can be quite lightweight. It also scales down better. The nuclear charges will be from 0.025 kilotons to 2.5 kilotons.

    The complicated stroke cycle is to smooth out the impulses from each blast, transforming it from a neck-braking jerk into a prolonged smooth acceleration.

    Jondale Solem calculates that the specific impulse is a function of the mass and yield of the nuclear charges, while the thrust is a function of the yield and explosion repetition rate. In this case, the mass of the nuclear charge is the mass of "propellant".

    Remarkably the mass of the spinnaker (sail) is independent of the size of its canopy or the number or length of its tethers. This means the canopy can be made very large (so the bomb blast radiation does not harm the canopy) and the tethers can be made very long (so the bomb blast radiation does not harm the crew). The mass of the spinnaker is directly proportional to the bomb yield and inversely proportional to the number of tethers.

    Medusa Sail Deceleration

         The program allows for the application of a Medusa Sail deceleration mechanism, using the theory as developed by Solem [14, 15, 16]. This involves a large sail canopy, connected by various spinnaker and servo-winches, to the main vehicle. Pellets or units are detonated inside the sail area, imparting a pressure force and thereby thrust in the opposite direction of motion. The working assumption in the current model is to use a high-strength polymer (e.g. polyethylene) which has a material density of around 990 kg/m3. The user also specifies the sail material Young's modulus of elasticity, tensile strength of the spinnaker material, which for the high-strength polymer are given values of 220 GNm2 and 5 GNm2 respectively. The distance to the detonation point from the spacecraft is also specified, as well as the time between detonations.

         The Specific impulse (Isp) of the Medusa Sail is given by:

         Where g is the acceleration due to gravity, Ap is the projected area of the canopy, r is the detonation distance from the spacecraft, E is the energy release per detonation, mb is the mass of an individual unit. This can then be multiplied by acceleration due to gravity to get the exhaust velocity vex:

         The impulsive pressure (P) delivered by each detonation is given by:

         Where t is the approximate debris expansion time per detonation.

         The average thrust (T) is given by:

         Where δt is the time between detonations.

         The approximate radius of the canopy debris cloud per detonation (rd) is given by:

         Where Eparticle is the approximate energy of the explosion per detonation, leading to the emitted particles.

         The mass of the sail canopy (mc) is given by:

         Where σmax is the tensile strength of sail material, Y is the Young's Modulus of elasticity, ρs is the density of sail material

    (ed note: for purposes of the study they looked at Ap=7.85e7 m2, r=1,000 m, mb=25 kg, E=100 GJ to 100,000 GJ, σmax=5 GNm2, Y=220 GNm2, and ρs=990 kg/m3)

    Medusa Exhaust Velocity (km/s): 1,042
    Medusa Specific Impulse (s): 106,220
    Total Medusa Unit Mass (tonnes): 750
    Radius of Gas Debris expansion from single unit (km): 20
    Maximum Conservative Spinnaker Mass (tonnes): 4
    Impulsive Pressure (N/m2): 2.48057942E-04
    Average Thrust (kN): 26,042
    Single Unit Mass (kg): 25
    Number Units: 30
    Wet (incl.Medusa) Mass (tonnes): 1,038.40002
    dry (incl.Medusa) Mass (tonnes): 1,033.65002
    Medusa delta V (km/s): 4.7759
    Pre-Medusa delta V (km/s): 36,985.7578      12.3371210
    Final effected velocity (km/s): 36,980.9805
    Percentage Medusa dv Reduction: 1.29167121E-02

         [14] Solem, J.C, The Moon and the Medusa: Use of Lunar Assets in Nuclear-Pulse-Propelled Space Travel, JBIS, 53, pp.362-370, 2000.
         [15] Solem, J.C, Nuclear Explosive Propulsion for Interplanetary Travel: Extension of the Medusa Concept for Higher Specific Impulse, JBIS, 47, pp.229-238, 1994.
         [16] Solem, J.C, Medusa: Nuclear Explosive Propulsion for Interplanetary Travel, JBIS, 46, pp.21-26, 1993.


    Inspired by a passing comment on the Eldraeverse Discord, we now present a galari starship, the Sapphire Coloratura-class polis yacht; the favored interplanetary and interstellar transport of all sophont rocks of wealth and taste.


    Operated by: Galari groups requiring luxurious private transit.
    Type: Executive polis yacht.
    Construction: Barycenter Yards, Galáré System

    Length: 96 m (not including spinnaker)
    Beam: 12 m (not including radiators)

    Gravity-well capable: No.
    Atmosphere-capable: No.

    Personnel: None required (craft is self-sophont). Can carry an effectively arbitrary number of infomorph passengers.

    Main Drive: Custom “dangle drive”; inertially-confined fusion pellets are detonated behind a leading spinnaker, the resulting thrust being transferred to the starship via a tether.
    Maneuvering Drive: High-thrust ACS powered by direct venting of fusion plasma from power reactors; auxiliary cold-gas thrusters.
    Propellant: Deuterium/helium-3 blend (pelletized aboard for main drive).
    Cruising (sustainable) thrust: 7.2 standard gravities
    Peak (unsustainable) thrust: 7.5 standard gravities
    Maximum velocity: 0.12 c (based on particle shielding)


    4 x galari body-crystals; since the galari are ergovores, any galari passenger or AI system may use these for EVA purposes.


    1 x standard navigational sensor suite, Barycenter Yards
    1 x lidar grid and high-sensitivity communications laser grid, Barycenter Yards


    Laser point-defense grid.

    Other Systems:

    • Cilmínár Spaceworks navigational kinetic barrier system
    • 4 x Bright Shadow secondary flight control systems
    • Kaloré Gravity Products type 1MP vector-control core
    • Systemic Integrated Technologies flux-pinned superthermal radiator system

    Small craft:

    5 x minipoleis (no independent drive systems; local accumulators only)


    The Sapphire Coloratura was intended to be a shining jewel in the crown of galari starship design, so it is perhaps fitting that it indeed resembles a shining jewel, the translucent crystal of its main body throwing sparkles of rainbow light everywhere when it chooses to fly close to stars, or when it is illuminated by the fiery blasts of its main drive.

    The main body of the ship is similar to, in many ways, the galari themselves; a sixteen-faceted crystal, with eight long facets facing forward to the bow tip, and short, blunter facets facing aft towards the mechanical section, a gleaming metal cylinder with a rounded-off end taking up the remaining two-thirds of the starship’s length.

    To proceed from fore to aft, the bow tip of the ship is capped with metal, housing the core mechanisms of the dangle drive; the sail deployment system, tether terminus, pellet launcher, and ignition lasers.

    From our Earth perspective, this drive is very similar to the Medusa-type Orion; thrust is delivered to the starship via a 216 m diameter spinnaker “sail” on a tether ahead of the craft. Rather than dedicated pulse units, the drive projects pelletized D-3He charges ahead of the craft to the focal point of the spinnaker, where inertially-confined fusion is initiated by the ignition lasers, reflected to surround the pellet by the inner surface of the spinnaker. The resulting nuclear-pulse detonation accelerates the craft, smoothed out by the stroke cycle of the tether (see above link).

    The main crystal body of the craft is essentially a solid-state piece – save for cooling labyrinths and the axial passage required by the drive – of galari thought-crystal: a substrate which holds the ship’s own intelligence, those of all passengers and any crew needed, along with whatever virtual realms, simulation spaces, or other computational matrices they may require. As such, there is little that can be described by way of an internal layout; most polis-yachts are unique in this respect.

    The “waist” – broadest point – of the body is girdled by a machinery ring, containing within it the four fusion power reactors (multiple small reactors were preferred for extra redundancy by the designer) with the associated ACS, and at points between them, the backup flight control systems, navigational sensor suite, and other small auxiliary machinery.

    At the aftmost point of the main body, where the blunter end of the crystal joins the mechanical section, eight crystal spikes project, symmetrically, from the point of junction. These are left hollow by the manufacturer and equipped with tip airlocks to provide a small amount of volume for cargo space and aftermarket customization; if non-ergovore passengers are expected, two of these are typically converted into quarters and life-support. A central chamber where the spikes meet serves as a body and robot hotel.

    Entering the mechanical section, an accessible chamber at the forward end of the cylinder provides accommodation for the vector-control core and larger auxiliary machinery, including the thermal control system. The remainder of the section is entirely made up of bunkerage for the reactors and main drive.

    The galari have never, it should be noted, shied away from making maximal use of vector control technology. This is particularly notable in the Sapphire Coloratura‘s design in two areas:

    First, its radiators, which cloak the center of the mechanical section with a divided cylinder of gridwork, individual carbon-foam emitting elements held together and in place away from the hull by vector-magnetic couples, linked back to the ship itself only by the ribbons of thermal superconductor transmitting waste heat to them; and

    Second, by the minipoleis that the Coloratura uses as small craft. Resembling nothing so much as miniature duplicates of the starship’s main body, these auxiliary blocks of thought-crystal are held in place orbiting the main body of the ship – often in complex patterns, even under full acceleration – connected only by vector-magnetic couples and whisker-laser communication.

    That is pure ostentation.

    Inertial Confinement

    Exhaust Velocity10,000,000 m/s
    Specific Impulse1,019,368 s
    Thrust100,000,000 N
    Thrust Power500.0 TW
    Mass Flow10 kg/s
    Total Engine Mass1,000,000 kg
    Specific Power2.00e-03 kg/MW

    A pellet of fusion fuel is bombarded on all sides by strong pulses from laser or particle accelerators. Some use a two dimensional ring of lasers like a proverbial circular firing squad. Others expand it into a three dimensional spherical firing squad. The beams implode the pellet, raising the density and temperature to the point where a fusion reaction ignites.

    The inertia of the fuel holds it together long enough for most of it to undergo fusion, instead of using a magnetic bottle as in Magnetic Confinement fusion.

    The spherical arrangement of lasers would have a gap in it for the exhaust nozzle.

    D-D Fusion Inertial
    Exhaust Velocity78,480 m/s
    Specific Impulse8,000 s
    Thrust3,200 N
    Thrust Power0.1 GW
    Mass Flow0.04 kg/s
    Total Engine Mass243,333 kg
    Frozen Flow eff.50%
    Thermal eff.50%
    Total eff.25%
    ReactorInertial Confinement
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorAblative Nozzle
    Specific Power1,938 kg/MW

    A “target” of fusion fuel can be brought to ignition by “inertial confinement”: the process of compressing and heating the fuel with beamed energy arriving from all sides. A snowflake of deuterium, the “heavy” isotope of hydrogen, can be imploded and fused with a combination of lasers and deuterium particle beams.

    The illustrated design uses combined input beam energy of 38 megajoules, arrayed in a ring surrounding the ejected iceball target. This energy operates at 1 Hz to blast a 2 gram ice pellet ejected each second. The outside 99% of the pellet is ablated away within 10 ns, super-compressing the deuterium fuel at the core to a density of a kilogram per cubic centimeter. The T and 3He products are catalyzed to undergo further fusion until all that remains is hydrogen, helium and some neutrons. (Neutrons comprise 36% of the reaction energy.) Fractional burn-up of the fuel (30%) is twice that of magnetic confinement systems, which implies a 40% higher fuel economy. The energy gain factor (Q) is 53.

    For a 500 MWth reactor, 320 MW of charged particles are produced, which can be used directly for thrust or metals refining. About 105 MW of fast neutrons escape to space, but another 75 MW of them are intercepted by the structure. About two thirds of this energy must be rejected as waste heat, but the remainder is thermally used to generate electricity or to breed tritium to be added to the fuel to facilitate the cat D-D pellet ignition.

    When used as a rocket, an ablative nozzle, made of nested layers of whisker graphite whose mass counts as propellant and shadow shield, is employed (much like the ACMF).

    “A Laser Fusion Rocket for Interplanetary Propulsion,” Hyde, R., 34th International Astronautical Conf., AIF Paper 83-396, Budapest, Hungary, Oct. 1983.

    (To keep radiator mass under control, I reduced the pellet repetition rate from 100 Hz to 1 Hz).

    From HIGH FRONTIER by Philip Eklund

    ENZMANN: There has been a twenty-year world-wide effort to tame the elusive geni of fusion (an apt analogy because during that time the primary line of research was directed toward trapping a plasma within a strong magnetic “bottle” until enough of it could “fuse” to release more energy than the containment process used). It was twenty years marked by dogged perseverance, brilliant theoretical insight, and agonizing frustration as physicists practically had to invent a whole new branch of Physics, something with the unpronounceable title — “Magnetohydrodynamics.”

    The essence of the problem was simple. For each new technical innovation devised to contain the plasma long enough for fusion to occur, the plasma devised three new techniques to escape. It was somewhat like trying to trap an angry anaconda with rubber bands.

    Into this dismal arena (for prospects of controlled fusion seemed to be beyond the end of the century, until recently), a new idea was injected. Quite simply, it asked, “Is there anything in the rules of this exotic game which insists that fusion has to be a continuous process? Suppose, instead of pursuing the exceedingly difficult road of containing the plasma indefinitely, with fusion energy trickling out, we examine an alternative technique of compressing the plasma intermittently. If we do this sequentially and often enough, the effect will be identical to a steady trickle of fusion power such as we have been laboriously pursuing with other techniques. The secret here is that all we have to do is hit that plasma hard enough, from all sides, and plain old inertia will do the rest. The result will be a fusion power plant somewhat similar to an internal combustion engine where it is the average power of the eight sequential explosions in the cylinders which moves the car. Of course, there is one problem: what are we going to hit the plasma with in the first place?”

    It was at this point that another technical breakthrough came to the rescue. Enter the LASER. It was suggested that intense light from a large laser, a source of precisely controllable electro-magnetic radiation, could be directed via suitable optics into a very small volume Of space. If, as the laser was fired, a tiny pellet of fuseable material (in this case, frozen deuterium-tritium) were dropped precisely into that tiny volume, the following would happen:

    The focused electro-magnetic radiation would exert a pressure on the pellet. Focused to such a small volume and arriving in such quantity, the prqssure would be enormous. But more important, the frozen deuterium (heavy isotopes of hydrogen, the most abundant element in the Universe) would absorb this energy at its surface which would, of course, be violently heated. It was Dr. John Nuckolls (of the Lawrence Radiation Laboratory, attached to the Univ. of Calif.) who predicted what would happen.

    If the laser pulse is actually two pulses—a smaller pulse of energy followed immediately by the main blast—then the outside of the pellet will absorb the first pulse, vaporize, and thus surround the pellet with a sort of atmosphere. The main pulse of intense laser light, upon arriving, will then further heat this “atmosphere,” causing it to expand rapidly in all directions. At this stage, out tiny pellet of frozen deuterium resembles very much a tiny synthetic star with a superheated atmosphere blowing off explosively in all directions. Now, Newton enters the picture, for the reaction pressure to this explosive departure of the shell of gas surrounding the pellet drives shockwaves from all sides into the tiny, icy heart of this miniature “star.” These crushing pressures actually compress the center of the pellet into the densities and temperatures found within the centers of stars like our sun, a million miles across! Yet the star created in the center of the vacuum chamber is only barely visible to the naked eye, a few tenths of a millimeter in diameter.

    At such densities and pressures the deuterium-tritium mixture fuses, the nuclei colliding and transmuting themselves into helium in a violent analogue of processes that take millions of years to complete in “normal-sized stars” such as our sun. The liberated fusion energy blasts apart the heart of this newly created “star,” sending the reactants in all directions at appreciable fractions of the speed of light. The tiny star explodes, ending its brief existence—as do many of the real stars a trillion times as large—as a supernova. Total time from creation to the death of this artificial star is less than a millionth of a second.

    If this process were accomplished successfully, the released energy would be greater than that required to fire the laser, maintain the vacuum, and keep the fuel frozen in precisely deliverable pellet quantities. Repetition of the entire sequence at rates ranging from once per second to a hundred times per second would provide for the “average” flow of energy envisioned previously.

    There is enough deuterium on Earth alone to provide essentially unlimited energy for the remaining lifetime of the sun. This is readily agreed upon by a wide spectrum of energy and environmental experts. However, what some may have apparently failed to consider, is that there are stores of deuterium within the rest of the solar system which dwarf our terrestrial supply (which would come primarily from seawater.) One of the moons of Jupiter, Callisto, for instance, which Pioneer 10 has revealed to have a density of only 1.65, seems to be a satellite with a 3,000 mile diameter, composed essentially of ice! In that ice, there is a vast supply of deuterium waiting to be mined. Little thought is necessary to realize that with fusion, which gives us the solar system through fusion rockets, we will discover enormous quantities of fuel beyond the “meager” quantity presently on Earth. No, energy is decidedly not a problem of the future.

    It naturally occurred to an awful lot of people at about the same time that if you aim a high-power laser through magnifying glass at a deuterium pellet, fusion will be in hand . Those same people were shortly seen rapidly retreating behind their office doors, biting their thumbs with a vacant look, pounding their heads against the nearest convenient hard object, or just standing staring down at their shoelaces. Word quickly spread (bad news may exceed light speed under certain circumstances) that what you achieved when you aimed a high-power laser at a deuterium pellet through a magnifying glass or any other lens system, was not a fusion reaction, but a mad scramble ofwhite-coated laboratory personnel for the cover of the near desks, benches, closets, and lavotories as the precision optics promptly exploded in about a million high temperature fragments, thus wiping out both the experiment and initial high hopes for fusion from lasers.

    The problem, of course, was discontinuities in the glass. At the energy densities emitted by the lasers in question, the slightest imperfection (and even the finest optical glass is not perfect) absorbs enough energy to flash instantly into vapor, shattering the lens. These lasers had been used in an M.I.T. project designed to blast tunnels through solid granite! They had shattered cobblestones into screaming shards of high velocity shrapnel and had punched through inch-thick destroyer plate. The power in these beams exceeds that of the sun by factors approaching the millions. No lens system could have the perfect transparency necessary to withstand even the shortest pulse of these hellish systems, not even for the billionths (10-9) or trillionths (10-12) of a second it would take to produce the deuterium “star.” Fusion, ignited by the “match” of the laser, seemed impossible to achieve for the very reasons which had first made it so attractive. How do you set off a reaction with a match you cannot hold?

    If the problem of the lens could be surmounted, Man would succeed in an incredible quest—bringing the stars, essentially, to Earth, harnessing their unlimited potential to solve Man’s most threatening problems for all time. But to succeed, the problem of the lens had to have its solution. The answer, as with all answers in the Universe to fundamental problems, was the picture of elegant simplicity: instead of attempting to focus the laser with a lens, use a mirror! With a suitable mirror system, the ravening power of one, ten, or a hundred laser systems could be spread out across a reflecting surface so that the power density per unit of area was well below the level required even to significantly raise the temperature of the mirror. Then, with appropriate geometry, this dilute radiation could be brought to a needlepoint focus on the deuterium target. Result: FUSION. Simple. Elegant. Practical.

    I like to think the first solution was the result of our work in Northeast Cryonics, but that isn’t quite correct. Perhaps, however, ours, I feel, is one of the least expensive and most elegant solutions. One of our designs, in fact, can be seen in Freff s illustration for this article.

    The secret of fusion-by-laser lies in discovering the correct geometry for the mirror system, a geometry which will first dilute the radiation from lasers of incredible power, lasers which are now being constructed or are within sight of our present state-of-the-art. With those two tools, creation in a fusion reactor of actual “stars” identical in every respect to those we see spangling the night sky will become a reality. And a new Age will have begun for mankind. That is the scientific breakthrough now only instants away by the standards of the cosmic clock.

    Examine carefully, if you will, the diagram I alluded to previously. You will notice that laser energy enters from two regions, is combined by appropriate optical flats, then flashes down into a conical mirror protected from destructive heating through a process identical to the reason we get cold in the winter and hot in the summer—grazing incidence of incoming energy. The ring of laser energy (since, of course, this system is three-dimensional) is reflected all along the mirror length, as you can see, and is brought to a tiny pinpoint focus of searing power at the precise special point where it will impact a pellet of deuterium mixture.

    Now, note very carefully. The end of this arrangement of optics, lasers, and hyperbolic mirror is open to space! And what is a container, closed on three sides, in which a high-temperature gas is expanding? Obviously, a rocket.

    This is why we shall have fusion torchships before we have fusion reactors lighting cities. You have to proceed through the stage of rocket development to get to power reactors.

    Spaceflight, up to now, has consisted of assembling a very large quantity of fuel and oxidizer in a metal tube, together with an engine to burn it in, and a very small payload (relatively speaking). The object: to burn said propellants as soon as possible, thus accelerating fuel, tank, engine and payload away from the earth as rapidly as possible. Once having released the payload from the final stage, you are in for the Long Wait.

    Since you have used most of the energy contained in the fuel you burned just to climb to the top of Earth’s gravitational “hill,” there isn’t much left to give you a healthy push on your way to Mars, the Moon, Venus or Jupiter. There you are, coasting along at a few thousand miles per hour, with millions of miles to cover. Worse, you are not traveling in a straight line, but a curving path which will add literally hundreds of millions of miles to the trip. This is the classic, coasting interplanetary journey portrayed by everyone from Asimov to Zelazny.

    Therefore, later—much later—varying from about a hundred days for a “quick” trip to Venus to almost two years to Jupiter (or six years to Saturn), you arrive. By which time most people will have forgotten you’re out there, except for that dedicated gang back at Mission Control.

    That is the picture of spaceflight held by 99.99 percent of those who take the trouble to think seriously about the subject at all.

    The days of such absurdities are essentially over. Alas, the 145-day flights to Venus and the 240-day missions to Mars (with its exciting layover of a Martian year, waiting for Earth to be in correct position for the return coasting flight) are not going to happen. Technology and time have caught up, even with Arthur Clarke (at least in the solar system) (alas, that turned out not to be the case).

    This is the way it will happen: A torchship will leave earth orbit, accelerating on a series of explosions produced by the events previously described. Fusion, with millions of times the energy of chemical reactions, and much higher coupling of this energy to the exhaust, results in a rocket engine and total ship performance which is nothing less than sensational, compared to the previous scenario. The ship accelerates away from Earth at (let us say, for the purpose of making a point), one “gravity”—an acceleration equivalent to that experienced on the surface of Earth where all objects fall at the same rate, 32 ft/sec2

    To those of us in the ship, everything will appear as it does on Earth, unless we look out the windows. Something dropped in the cabin will apparently “fall” to the floor at the same rate as it would on Earth, although actually it is the floor which is accelerating up to meet the object. No difference. Our “weight,” synthetic as it is and produced only by the thrust of our fusion engine, will be just as useful when it comes to sitting in chairs, or walking, or drinking liquids as actual gravitational weight is on Earth.

    After about half an hour of this, we may begin to wonder if there isn’t something rather strange about the performance of our vehicle. After all, we know that “normal” rockets such as those used in Apollo never burned for longer than about 12 minutes, and in that time consumed literally thousands of tons of fuel! This ship, about as massive as a jumbo jet (200 tons) has, according to our fuel gauge, consumed only 1800 pounds of deuterium—less than a ton! And we are still accelerating.

    After about two hours of this novel experience, enjoying normal earth weight, carpeted lounges, the pleasant attentions of stewards who bring drinks we can consume in the normal manner, we walk down to the observation area where we see a magnificent panorama of space. Imagine the shock as we realize that we are halfway between Earth and the moon—120,000 miles away from Earth and still accelerating, still consuming fuel. At this point we discover something else. The heavens are slowly turning around; the earth which was behind us is now in front, and the moon which was before us has taken up a position aft (to borrow a nautical phrase). It becomes apparent to us almost at once that the two celestial objects have not changed positions at all. Our 200-ton ship has rotated 180 degrees (Heinlein calls this a "skew flip", The Expanse calls it a "flip-and-burn"). Our fusion “torch” engine, which only moments before was accelerating us to higher and higher velocities, is continuing to thrust, but its effect is now to kill our enormous velocity. Inside, of course, nothing changes. Acceleration, deceleration, it’s all the same within the ship where the floor still presses reassuringly against our feet with normal Earth weight.

    Glancing at the system read-outs positioned so conveniently near the view-glass, we are incredulous at our measured velocity indicator. It must be in error! We are moving, it says, at 63 kilometers per second, relative to Earth. We remember Apollo, when the Command Module crawled across this point at a mere .6 kilometers per second—one hundredth our velocity—not in two hours, but over two days! Two hours more and the Captain slides us neatly into a breathtaking orbit of the moon. Under its gravitational field, weightless for the first time during the flight (the drive is turned off), we take 45 minutes to swing around the Farside and head home. Earth before us once again, the torch is lit, normal gravity resumes, and we accelerate away from Luna. It has taken us less time to span the almost half million miles between these two worlds than it took, in 1974, to fly across the United States in a vehicle of comparable size. It has taken far less fuel (a total of about 15 tons of deuterium-tritium) than a 747 burned flying cross-country (about 70 tons of kerosene) and at vastly higher peak velocities—227,000 km/hr, compared to a 747’s 960 km/hr. We have made the trip totally contained in an essentially terrestrial environment with all the “comforts of home.”

    Similar journeys, to any planet in the solar system, under identical environmental conditions (simulated Earth gravity through continuous acceleration /deceleration; conventional meals served on real plates; and beverages served in cups or glasses) will be commonplace within 15 years. No point in the solar system will be further from any other point, at one “G” acceleration, than 5 weeks’ travel time, the length of a relaxing terrestrial cruise. The ships, of course, will resemble ocean liners far more than they will terrestrial aircraft

    From TORCHSHIPS NOW! by Robert D. Enzmann and Richard C. Hoagland (1974)
    Propulsion SystemIC Fusion
    Exhaust Velocity170,000 m/s
    Specific Impulse17,329 s
    Thrust240,000 N
    Thrust Power20.4 GW
    Mass Flow1 kg/s
    ReactorInertial Confinement
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Wet Mass6,000,000 kg
    Dry Mass1,835,000 kg
    Mass Ratio3.27 m/s
    Width170 m
    Height100 m

    Magneto Inertial Fusion

    Magneto Inertial Fusion
    Specific Impulse50,000 s to 100,000 s
    Gear Shift Isp5,000 s to max
    Specific Power10 kW/kg to 100 kW/kg

    There are two main approaches to utilizing nuclear fusion, magnetic confinement and inertial confinement. Magnetic confinement uses titanic magnetic fields, inertial confinement is how fusion bombs explode (a third way would be stars shining by gravitational confinement, but we don't know how to generate artificial gravitational fields).

    Inertial confinement ignites the fusion fuel by imploding a solid pellet of fuel with a circular firing squad of lasers or particle beams. This raises the density and temperature high enough for fusion ignition. It confines the burning fusion fuel by sheer inertia. That is, it is hoping that the burning fuel simply does not have enough time to expand below fusion density before all the fuel is burnt.

    Magnetic confinement ignites using a magnetic field to squeeze a cloud of fusion fuel plasma until it is hot and dense enough to ignite. It confines the burning fusion fuel with the same magnetic field. More like tries to confine, the blasted plasma keeps wiggling out of the cracks in the magnetic field before it is all burnt.

    As propulsion systems, both have major drawbacks.

    Problem 1

    Magnetic confinement requires huge (read: massive) electromagnets. The technique also has the problem of plasma instabilities (read: fusion plasma has thousands of different ways to wiggle out of the magnetic cage) which so far have defied any solution. Meaning that every time fusion researchers have devised a new magnetic cage, the blasted plasma finds two new ways of wiggling out.

    Inertial confinement works well in bombs, but trying to do it in a small controlled fashion (read: so the fusion reaction does not vaporize everything in a one kilometer radius) has also defied any solution. The compressing laser or particle beams have such low efficiencies that tons of excess power is required. Timing all the beams so they strike at the same instant is a challenge.

    Problem 2

    Both approaches have a problem with getting the fusion reaction energy to heat the propellant. Magnetic confinement tries to use the actual fusion plasma as propellant, resulting in a ridiculously small mass flow and thus a tiny thrust.

    Problem 3

    Also, there is nothing in between the fusion reaction and the chamber walls, leading to severe damage to the walls. The escaping radiation harms the crew as well.

    The proposed solution is to combine magnetic confinement and inertial confinement.

    This is called Magneto Inertial Fusion (MIF) or Magnetized Target Fusion (MTF). For some researchers there are subtle differences between what mechanism the two terms describe, for other researchers the two terms are interchangeable. I'm going to use the terms as synonyms because the subtle differences are too subtle for my tired brain to distinguish.

    A blob of fusion fuel is converted into plasma. It is captured in a field reversed configuration. This means the plasma is shaped into a spinning torus, much like a smoke ring. If a smoke ring had a temperature of hundreds of thousands of degrees. The smoke ring is stablized by a magnetic field, that is, it is held by magnetic confinement.

    The point is the smoke ring of plasma is self-stable, once its internal magnetic field is established the ring does not need external electromagnets to prevent it from dissipating like a fart in a windstorm. What the external electromagnets does do is constrain the ring to travel down the axis of the engine. The smoke rings magnetic field is reversed with respect to the external magnets, so it is repelled by them (that's why it is called field reversed configuration). The repulsion prevents the smoke ring from collding with the engine walls.

    The smoke ring travels down the axis until it enters the reaction chamber.

    In the chamber, the smoke ring of fusion fuel is brutally crushed by a "liner", causing fusion. In standard inertial confinement fusion the fusion pellet is crushed by a circular (actually spherical) firing squad of laser beams or particle beams. But here the fuel is crushed by a liner made of matter. The liner is crushed by a magnetic field, but a much weaker field than is required by magnetic confinement fusion. Which is generated by electromagnetic coils which are much lower in mass than the ones used in MC fusion.

    In some designs the liner is a ring of metal foil. In others the liner is a circular firing squad of plasma jets.

    And in others, they do not bother with the FRC smoke ring of fusion fuel. Instead, the circular firing squad of plasma jets first fires a tiny jet of fusion fuel, immediately followed by a longer jet of plasma to use as a liner.

    What advantages does this Rube-Goldberg contraption give us?

    Trying to compress fusion fuel with immaterial laser beams is inefficient. Try hammering a nail using the beam of light from your flashlight if you don't believe me. Using a liner made of metal is vastly more efficient, which is why we pound nails using hammers made of matter. This makes MIF better than inertial confinement fusion. Attempting to compress using immaterial magnetic field is worse, all the problems of laser beams plus plasma instabilities on top of that. This makes MIF better than magnetic confinement fusion. In other words: Problem 1 solved.

    Since the fusion fuel is being crushed by the liner, pretty much all of the exploding fusion energy released is going to be absorbed by the liner. For "liner" read "propellant." So unlike inertial confinement and magnetic confinement engines, the MIF will be very efficient at transfering the fusion energy into the propellant where it should be. Problem 2 solved.

    And the more of the fusion energy intercepted by the liner/propellant, the less damaging blast hitting the reaction chamber walls and the less deadly radiation inflicted on the crew. Problem 3 solved.

    In addition, the lack of gigantic magnetic coils and batteries of laser beams means the MIF engines have much lower mass than magnetic confinement and inertial confinement engines.

    Thio MTF

    Thio Magnetized Target Fusion
    Fusion Yield per pulse<1 GJ
    Fusion Gain70
    Specific Impulse77,000 sec
    Exhaust Velocity760,000 m/s
    Engine Mass41,000 kg
    Pulse Rate 40 Hz
    Jet Power25 GW
    Thrust66,000 N
    Specific Power α605 kW/kg
    Specific Thrust1.5 N/kg
    Pulse Rate 200 Hz
    Jet Power128 GW
    Thrust340,000 N
    Specific Power α1,141 kW/kg
    Specific Thrust3 N/kg

    This is from High-Energy Space Propulsion based on Magnetized Target Fusion (1999)

    This is a Magneto Inertial Fusion concept where the liner is jets of plasma.

    They calculate that the hydrogen liner plasma is capable of converting more than 97% of the otherwise wasted neutron flux into charged particle energy available for thrust. Which is fantastic since otherwise 38% of the fusion energy is wasted. This also allows the heavy radiation shielding to be reduced, since neutrons are very damaging to engines and deadly to the crew.

    Theoretically it is possible to operate a magnetic nozzle in such a manner that it would have a nozzle efficiency of 80% in converting the spherically radial momentum of the exploding fusion into axial rocket thrust AND also harvesting some of the energy as electrical power. This is done with a magnetic flux compression generator. As per standard operating procedure, this is used using the camera strobe principle, where energy is harvested from one pulse and used to ignite the following pulse.

    A plus rate of up to 200 Hz appears to be practical, given the simplicity of the electrical circuit.

    Unlike most inertial confinement fusion engine, the MTF does not require intricately machined and layered fuel pellets. Just a tank of deuterium fuel and you are good to go.

    The Thio MTF had a very low system mass and volume, high thrust and Isp, high fusion gain, relatively low thermal waste, and the aforementioned ability to convert 97% of the neutron flux into thrust.

    Just like the HOPE MTF a pair of opposed conical theta pinch guns fire the plasma smoke-rings which collide at the target location to created the FRC magnetized target plasma. This is imploded by a barage of 32 plasma accelerators, whose plasma jets merge to form a spherically converging plasma liner. The implosion heats the target plasma to thermonuclear fusion temperature, and it undergoes fusion.

    Now, also just like the HOPE MTF, the plasma liner jets are actually in two parts. Each jet starts as a small spurt of deuterium gas then becomes a long jet of ordinary hydrogen. So the spherical plasma liner has a thin inner layer of deuterium fusion fuel, and a think outer layer of hydrogen propellant.

    The thermonuclear explosion makes the plasma liner think it hit a brick wall. The inner deuterium layer of the plasma liner is compressed to extremely high density, and undergoes fusion. The fusion of the target plasma is just a spark, the fusion of the deuterium layer is the main event.

    The outer hydrogen layer is heated by the deuterium fusion, including 97% of the neutrons. This become the superheated propellant. The lower half of the propellant shoots out of the magnetic nozzle, creating thrust. The upper half of the propellant slams into the magnetic nozzle, is reflected downward, and also creates thrust. The magnetic nozzle is created out of magnetic lines of force generated by cryogenically cooled electromagnetic coils ("thrust coils").

    The thrust coils also contain the magnetic flux compression generator, which steals a bit of thrust energy and converts it into electricity. This is stored in the capacitor bank, and used in the next pulse.

    Also, like many pulse propulsion systems, a large shock absorber will be needed between the engine and the payload to prevent the spacecraft from being jolted to pieces. This smooths out the impulse.

    A system of heat radiators is used to dispose of waste heat; which comes mostly from the plasma guns, the electrical recharge system, and neutron heating.

    On top will be the customary anti-radiation shadow shield to protect the crew.

    HOPE (MTF)

    HOPE (MTF)
    Inner Liner (fuel)deuterium
    Outer Liner (propellant)hydrogen
    Specific Impulse70,485 s
    691,460 m/s
    Thrust5,798 N
    Thrust Power2.038 GW
    Waste Heat
    492.9 MW
    MTP Pulse Rate20 Hz
    Engine Mass121,333 kg
    Crew Radiation
    <0.05 Sv/yr
    Mass Breakdown
    Med-Temp Radiators22,340 kg
    Hi-Temp Radiators20,523 kg
    THERMAL TOTAL51,391 kg
    MTF Engine
    Mass Breakdown
    x49 Plasma Guns1,167 kg
    Capacitors3,502 kg
    x2 Theta Pinch350 kg
    20,576 kg
    Neutron Shield
    12,551 kg
    Nozzle Coils35,000 kg
    energy storage
    3,000 kg
    Recharge Circuit1,664 kg
    Neutron Shield
    (water tank)
    37,000 kg
    Power Cables2,000 kg
    ENGINE TOTAL116,021 kg

    This is from Conceptual Design of In-Space Vehicles for Human Exploration of the Outer Planets (2003).

    This is a Magneto Inertial Fusion concept where the liner is jets of plasma.

    For the rest of the ship, go here.

    Two theta (Θ) pinch guns fire magnetized blobs of easily ignitable D-T fusion fuel plasma on a collision course. They collide at the parabolic focus of the magnetic nozzle. The nozzle is a magnetic field formed by the thrust coils, because a nozzle made out of matter would be damaged by the fusion explosion.

    A split second behind the firing of the theta pinch guns, a battery of 48 plasma guns fire plasma jets targeted at the fusion fuel blob. These jets form a spherical "liner" around the fusion fuel.

    The jets are actually composed of two different plasmas. The inner bit is fusable deuterium fuel, the large outer part of the jet is hydrogen propellant.

    The liner collapses at about 750 kilometers per second and squeeze the fusion fuel like a nutcracker from hell. The fusion fuel ignites like a miniature H-bomb, which it is.

    The important points are that D-T fusion is easy to ignite (Lawson criterion of only one), but the reaction emits a relatively large amount of damaging neturons (79%) and uses expensive tritium. D-D fusion is much harder to ignite (Lawson criterion of 30), but the reaction only emits 38% neutrons and deuterium is very cheap. The idea is that the liner will ignite the tiny D-T fuel blob in the center, and the fusion explosion will be enough to ignite the huge amount of D-D fuel in the liner.

    The thermonuclear detonation drives the hydrogen propellant in the liner outwards, where it rebounds off the magnetic nozzle, producing thrust on the thrust coils, which produce thrust on the structural tapered spines, which produce thrust on the thrust frame of the spacecraft's spine.

    A large tank of water perched on top of the magnetic nozzle acts as an anti-radiation shadow shield, to protect the crew.

    As is standard operating procedure with many such pulse engines, some of the energy of the detonation is tapped and stored in capacitors. This energy is used to power the next pulse (for the plasma guns and to create the magnetic nozzle). Electrical current is induced in the coils as the plasma cloud expands. Each plasma gun has its own capacitor to store power for the next pulse. The energy for the magnetic nozzle is stored in something called a Superconducting magnetic energy storage (SMES) device, located just below the nuclear reactor.

    For the first pulse each of the capacitors and the SMES has to be slowly charged up by the nuclear reactor (since the poor little one-lung SP-100 can only crank out a pathetic 300 kilowatts). With subsequent pulses the capactors are recharged almost instantly, by the power of nuclear fusion.

    Slough FDR

    Slough Fusion Drive Rocket
    Exhaust Velocity50,420 m/s
    Specific Impulse5,140 s
    Remass AccelThermal Accel:
    Reaction Heat
    Low Gear
    Thrust103 N
    Thrust Power2.6 MW
    Mass Flow2.00×10-03 kg/s
    Delay between
    Fusion Pulses
    180 seconds
    High Gear
    Thrust13,800 N
    Thrust Power0.3 GW
    Mass Flow0.27 kg/s
    Delay between
    Fusion Pulses
    14 seconds

    Dr. John Slough and his associates have come up with a new type of magneto inertial fusion propulsion. You can find their published papers on the subject here (2012)

    This is a Magneto Inertial Fusion concept where the liner is a ring of metallic foil.

    In their design, the liner is a foil ring composed of lithium, about 0.2 meters in radius. Each liner will have a mass of 0.28 kg (minimum) to 0.41 kg.

    As the liner travels axially down the chamber, electromagnets crush it down into a solid cylinder (the crush speed is about 3 kilometers per second, the cylinder will have a radius of 5 centimeters). This is timed so that the plasma blob (plasmoid) is in the center of the cylinder. The liner compresses the plasmoid and ignites the fusion reaction.

    The fusion reaction vaporizes the lithium liner. The ionized lithium (plus the burnt fusion fuel) exits through a magnetic nozzle, providing thrust.

    In other words it both ignites and confines the fusion fuel with a collapsing wall of solid metal. The metal is being squeezed by an external magnetic field even as the fusion reaction is raging, which does a much better job of confinent than simple inertia or a rubbery magnetic field.

    Since this is an open-cycle system, the exhaust acts as the heat radiator, so the spacecraft can get by with only a tiny radiator. The energy to run the magnets can be supplied by solar cell arrays. Since the compression is so efficient, this will work with several types of fusion fuel: D-T, D-D, and D-3He. D-D is probably preferred, since tritium is radioactive with a short half-life, and 3He is rare.

    Please note that if you replace the magnetic nozzle with a magnetohydrodynamic (MHD) generator, the propulsion system is transformed into an electrical power generator. This could be used for ground based fusion power generators.

    Dr. Slough et al worked up two spacecraft for a Mars mission. The first was optimized to have a high payload mass fraction. The second was optimized to have the fastest transit time. Both were capable of a direct abort and return. The "Low Gear" engine is the study author's opinion of an engine easily achievable with current technology (that is, achievable fusion yields). The "High Gear" engine is a bit more speculative, but requiring only modest incremental improvements in technology.

    Fusion Drive Rockets (FDR)
    High Mass Fraction
    EngineLow Gear
    Transit Time90 days
    Initial Mass90 mT
    Payload Mass Fraction65%
    Specific Mass4.3 kg/kW
    Shortest Transit Time
    EngineHigh Gear
    Transit Time30 days
    Initial Mass153 mT
    Payload Mass Fraction36%
    Specific Mass0.38 kg/kW

    Plasma Jet MIF

    Plasma Jet Magnetio Inertial
    Base Parameters
    Mass of
    plasma (g)
    Efficiency of rail
    & θ-pinch guns
    Initial jet
    velocity (km/s)
    Heat fraction
    for 2nd power
    Frequency (Hz)
    Target ΔV (c)0.08
    Target Burn
    Time (years)
    Resulting Ship Parameters
    Fuel Mass (t)55,50337,84325,228
    Velocity (km/s)
    Impulse (s)
    Thrust (MN)0.521.341.78
    Fuel Mass (N/kg)
    Jet Power (GW)311.432985.457961.07
    Alpha (MW/kg)0.00560.07890.3156
    Heat (GW)
    Mass (t)

    This is from Project Icarus: Analysis of Plasma jet driven Magneto-Inertial Fusion as potential primary propulsion driver for the Icarus probe (2013).

    This is a Magneto Inertial Fusion concept where the liner is jets of plasma.

    A blob of fusion fuel plasma is injected into the center of the reaction chamber. It is bombarded by a spherical firing squad much like classic inertial confinement fusion. The difference is:

    1. The fusion fuel is a blob of plasma, not a solid pellet.
    2. The fusion fuel plasma blob is magnetized.
    3. The firing squad does not fire lasers or particle beams. Instead it fires cylindrical jets of plasma. The plasma is made from some element with a high atomic weight, so it has some serious momentum and inertia

    In the table there are three columns for three estimates of the performance of an actual engine. These are labeled CON (Conservative), MED (Medium), and OPT (Optimistic). The report notes that the Medium column is probably good enough for an unmanned interstellar probe. The Conservative column is probably good enough for missions within the solar system.

    • Mass Of Plasma: mass of the fusion fuel blob
    • Efficiency of rail & θ-pinch guns: efficiency of the railguns shooting the plasma liner jets and the theta-pinch guns creating the fusion fuel blob
    • Initial Jet Velocity: how fast the plasma liner jets are imploding
    • Heat Fraction for 2nd Power: fraction of the total rejected heat that is being used for the secondary power needs of the spacecraft. Meaning that some of the waste heat will be sent through a generator to make power for the ship's avionics and whatnot
    • Firing Frequency: how many fusion detonations are ignited per second
    • Fusion Gain: how many times bigger is the fusion energy compared to the input energy. The return on your investment, in other words
    • Target ΔV: the delta-V requirements the report assumes will be needed by the proposed interstellar mission
    • Target Burn Time: the burn time requirements the report assumes will be needed by the proposed interstellar mission
    • Nozzle Efficency: efficiency of the magnetic nozzle
    • Fuel Mass: The total mass of fuel needed for the mission. This is also a first approximation of the ship's wet mass, since with such outrageous delta-V requirements the fuel mass will dominate the total mass
    • Exhaust Velocity: what it says
    • Specific Impulse: what it says
    • Thrust: what it says
    • Thrust/Fuel Mass: Thrust to total fuel mass ratio, which is pretty darn close to thrust to mass ratio
    • Jet Power: what it says
    • Alpha: power to mass ratio
    • Waste Heat: amount of the power that turns up as waste heat and must be gotten rid of before the ship melts
    • Radiator Mass: mass of the heat radiators required to cope with the waste heat

    As the jets converge on the fuel at 750 kilometers per second they merge to form a spherical "liner". The liner collapses, squeezing the fusion fuel like a nutcracker from hell.

    Meanwhile as the fusion fuel is squeezed, so is its magnetic field. The density of the magnetic field increases to a point where is makes a conventional magnetic-confinement fusion engine look anemic.

    The shock where the imploding liner contacts the surface of the fuel blob heats it up. The liner also compresses the fuel blob, and soon fusion will be ignited. The internal magnetic field helps keep it confined long enough to burn all the fuel.

    The exploding fusion blob hits the magnetic nozzle, compressing the nozzle's magnetic field. This acts like a trampoline, making the fusion plasma rebound out the exhaust nozzle, creating thrust. Which is the purpose of all rocket engines. Meanwhile some of the energy in the nozzle field compression can be harvested to charge up the capacitors for the next round.

    The main advantage this propulsion system has over inertial or magnetic confinement is a drastically lower power requirement. Lasers, particle beam accelerators, or giant magnetics are power hogs. In this system the liner plasma jets can be lauched with relatively low powered rail-guns. This means you do not need tons and tons of capacitors to hold the huge jolts of electricity the other systems demand.

    Antimatter Bottle

    This section has been moved here

    Antimatter catalyzed

    Nuclear fission pulse drives like Orion scale up well, since it is relatively easy to design a bigger bomb than the last one. However, physics seem to prevent the creation of a nuclear device with a yield smaller than about 1/100 kiloton (10 tons, 42 GJ) and a fissionable material mass under 25 kilograms. This is due to critical mass restraints.

    However, if a tiny sub-critical bit of fissionable material is bombarded by a few antiprotons, it will indeed create a tiny nuclear explosion. The antiprotons annihilate protons in uranium atoms, the energy release splits the atoms, creating a shower of neutrons, and a normal chain reaction ensues. Using antiprotons, yields smaller than 1/100 kiloton can be achieved. This can be used to create Antimatter catalyzed nuclear pulse propulsion


    Exhaust Velocity598,000 m/s
    Specific Impulse60,958 s
    Thrust55 N
    Thrust Power16.4 MW
    Mass Flow1.00e-04 kg/s
    ReactorAntimatter Catalyzed
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle

    Antiproton-initiated Microfusion. Inertial Confinement Fusion. See here.


    Propulsion SystemACMF
    Exhaust Velocity132,435 m/s
    Specific Impulse13,500 s
    Thrust180,000 N
    Thrust Power11.9 GW
    Mass Flow1 kg/s
    Total Engine Mass27,000 kg
    Uranium 235
    ReactorAntimatter Catalyzed
    RemassSilicon Carbide
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorAblative Nozzle
    Wet Mass707,000 kg
    Dry Mass345,000 kg
    Mass Ratio2.05 m/s
    ΔV95,020 m/s
    Specific Power2 kg/MW

    Antiproton-catalyzed microfission, inertial confinement fission. See here.

    Fuel pellets have 3.0 grams of nuclear fuel (molar ratio of 9:1 of Deuterium:Uranium 235) coated with a spherical shell of 200 grams of lead. The lead shell is to convert the high energy radiation into a form more suited to be absorbed by the propellant. Each pellet produces 302 gigajoules of energy (about 72 tons of TNT) and are fired off at a rate of 1 Hz (one per second). The pellet explodes when it is struck by a beam containing about 1×1011 antiprotons.

    A sector of a spherical shell of 4 meters radius is centered on the pellet detonation point. The shell is the solid propellant, silicon carbide (SiC), ablative propellant. The missing part of the shell constitutes the exhaust nozzle. Each fuel pellet detonation vaporizes 0.8 kilograms of propellant from the interior of the shell, which shoots out the exhaust port at 132,000 meters per second. This produces a thrust of 106,000 newtons.

    The Penn State ICAN-II spacecraft was to have an ACMF engine, a delta-V capacity of 100,000 m/s, and a dry mass of 345 metric tons. The delta-V and exhaust velocity implied a mass ratio of 2.05. The dry mass and the mass ratio implied that the silicon carbide propellant shell has a mass of 362 metric tons. The wet mass and the thrust implied an acceleration of 0.15 m/s2 or about 0.015g. It can boost to a velocity of 25 km/sec in about three days. At 0.8 kilograms propellant ablated per fuel pellet, it would require about 453,000 pellets to ablat the entire propellant shell.

    It carries 65 nanograms of antiprotons in the storage ring. At about 7×1014 antiprotons per nanogram, and 1×1011 antiprotons needed to ignite one fuel pellet, that's enough to ignite about 453,000 fuel pellets.

    The system is very similar to Positron Ablative.

    H-B inertial catalzyed fusion
    H-B cat inertial
    Exhaust Velocity156,960 m/s
    Specific Impulse16,000 s
    Thrust4,700 N
    Thrust Power0.4 GW
    Mass Flow0.03 kg/s
    Total Engine Mass65,089 kg
    Frozen Flow eff.86%
    Thermal eff.85%
    Total eff.73%
    ReactorAntimatter Catalyzed
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorAblative Nozzle
    Specific Power176 kg/MW

    The fusion of hydrogen and boron 11 is a clean reaction, releasing only 300 keV alpha particles, which can be magnetically directed. However, the H-B fusion will not proceed at temperatures less than 300 keV unless catalyzed using exotic particles.

    One possibility: replace the electrons in H-B atoms with stable massive leptons such as magnetic monopoles or fractionally-charged particles (the existence of these is hypothetical). The resulting exotic atoms can fuse at “cold” temperatures, allowing the exotic catalysts to be recycled.

    A second possibility is to use antiproton-catalyzed microfission to initiate the H-B fusion. If a hundred billion antiprotons at 1.2 MeV in a 2 nsec pulse are shot at a target of three grams of HB: 235U in a 9:1 molar ratio, the uranium microfission initiates H-B and releases 20 GJ of energy. Operating at a fifth of a hertz, hydrogen and boron 11 reacting at a rate of 145 mg/shot produces 2000 MWth. A shell of 200g of lead about the target thermalizes the plasma from 35 keV average to 1 keV, low enough that this radiation can be optimally transferred to thrust using a magnetic or ablative nozzle at 73% efficiency. The ejected mass per shot is 2.4 kg. The exotic catalysts are recycled. Catalyzed fusion enjoys an excellent thermal efficiency (86%) and thus a good thrust/weight ratio (3.2 milli-g), making it one of the best engines in the game. The specific impulse ranges between 8 and 16 ksec, depending whether spin-polarized free radicals are used as the hydrogen fuel.

    “Antiproton-Catalyzed Microfission/Fusion Propulsion Systems for Exploration of the Outer Solar System and Beyond”, G. Gaidos, et al., Pennsylvania State University, 1998.

    (I used the ICAN-II spacecraft design, modified from cat D-T to cat H-B fuel, and scaled way down from 1 Hz to 0.2 Hz, and 302 GW to 2 GW.)

    From High Frontier by Philip Eklund

    Radioisotope Positron

    Radioisotope Positron Propulsion
    Positrons / pulse1.96×1011
    79Kr source area200 cm2
    (8 cm radius)
    Positron plus rate0.006Hz to 80kHz
    D fuel density1×1030/m3
    Ignition burn depth250 nm
    Specific Impulse3×105 sec
    Exhaust Velocity2,943,000 m/s
    Initial 79Kr1 μg
    Maximum 79Kr14 g
    79Kr enrichment0.99
    Accumulator lifetime100 sec
    Thrust efficiency0.65
    Thrust10nN to 0.2N
    Maximum engine power2.1 MW
    ΔV60,000 m/s
    Mass Ratio1.0 !!!

    This is from Radioisotope Positron Propulsion NIAC Phase I Report (2019)

    As with all non-torchship engines, this propulsion system suffers from the "can only increase the exhaust velocity at the expense of the thrust" problem. But this is one of the extreme cases. The exhaust velocity is a whopping 2,943,000 meters per second. Meanwhile the thrust is a totally minuscule 0.00000002 to 0.2 Newtons.

    Having said that, the delta-V is a jaw-dropping 60,000 meters per second WITH A MASS RATIO OF 1.0. With many other propulsion systems, rocket designers are happy if the spacecraft is only 75% propellant and 25% everything else. A spacecraft with Radioisotope Positron Propulsion is pretty much 100% rocket and payload, the propellant is only a few micrograms.

    Granted that a one metric ton space probe with such an engine will have an measly acceleration of 0.0001 meters per second (0.0125 snail-power), but you can't have everything. Be that as it may, the report compares their positron engine with an electric propulsion engine for a hypothetical capture/redirect of asteroid 2009BD and the positron engine kicks the electric engine to the curb.

    Theoretically you can use multiple engine arrays if you must have a higher thrust.

    The secret is positrons, i.e., antimatter electrons.

    Antimatter is the ultimate fuel, but trying to actually use the stuff has many practical problems. Manufacturing large amounts of antimatter is colossally expensive. And trying to safely contain that dangerous crap is like juggling live thermonuclear warheads during an earthquake.

    But the study authors realized that positrons occur naturally. Specifically there are several commonly available radioisotopes that emit positirons, such as sodium-22, cobalt-58, and krypton-79. By utilizing long-lived radioisotopes that emit positrons, one is storing the positrons safely inside the nuclei instead of trying to store them in a Penning Trap or something equally hazardous.

    Krypton-79 was selected as the fuel of choice. Each nuclei will eventually spit out a positron and decay into Bromine-79. It may be possible to convert the bromine into tetrabromomethane (CBr4) which can be utilized as a trapping or cooling gas. But the study did not look deeply into that.

    What's more, krypton-79 can be created during the mission (a "breeder" fuel cycle). The antiprotons will be used to catalyze deuterium-deuterium fusion, which creates lots of neutrons as a side-effect. These neutrons can be trapped by a layer of more-or-less stable krypton-78. The neutrons will transform krypton-78 into positron-emitting krypton-79. The 79Kr can be skimmed out and added to the fuel supply.

    The breeder layer of 78Kr will have to be pressurized to about 10 to 100 atmospheres, and be about one meter thick.

    Krypton-79 fuel gas will be frozen as a thin layer on a cryogenic surface. The positrons that escape will pass through a moderator to even out their energies. They then pass through the beam system. This transforms the continuously emitted positrons from a large area source into short pulses focused on a small spot size.

    The deuterium fuel is catalysed to form dense clusters. These are deposited on a moving tape substrate. The tape carries the deuterium clusters right into the small spot size of positrons. The annihilation of positrons with electrons create gamma rays. Each ray can kick a deuterium ion hard enough so it slams into another deuteron fast enough to undergo inertial confinement fusion. The charged particle fusion reaction products are ejected through a magnetic nozzle to create thrust. The neutron fusion reaction product are hopefully mostly captured by the krypton-78 blanket. The rest either escape into space or do terrible things to the spacecraft.

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