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.
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 NTR|
|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: 1I/'Oumuamua was that interstellar object that came streaking through the solar system in 2017. Everybody who was well read in science fiction novels were immediately reminded of Arthur C. Clarke's Rama. While the object is probably not an alien generation starship, you never know for sure until you examine the blasted thing. A pity it is traveling at 26.33 kilometers per second, we'll never catch it now.
Or can we?
The authors of the paper figured that using chemical rockets made about as much sense as using a arthritic tortoise to catch a cheetah. Nuclear thermal is more reasonable. They crunched the numbers for a variety of solid-core nuclear thermal rockets)
The first definite interstellar object observed in our solar system was discovered in October of 2017 and was subsequently designated 1I/’Oumuamua. In addition to its extrasolar origin, observations and analysis of this object indicate some unusual features which can only be explained by in-situ exploration. For this purpose, various spacecraft intercept missions have been proposed. Their propulsion schemes have been chemical, exploiting a Jupiter and Solar Oberth Maneuver (mission duration of 22 years) and also using Earth-based lasers to propel laser sails (1-2 years), both with launch dates in 2030. For the former, mission durations are quite prolonged and for the latter, the necessary laser infrastructure may not be in place by 2030. In this study Nuclear Thermal Propulsion (NTP) is examined which has yet to materialise as far as real missions are concerned, but due to its research and development in the NASA Rover/NERVA programs, actually has a higher TRL than laser propulsion. Various solid reactor core options are studied, using either engines directly derived from the NASA programs, or more advanced options, like a proposed particle bed NTP system. With specific impulses at least twice those of chemical rockets, NTP opens the opportunity for much higher ΔV budgets, allowing simpler and more direct, time-saving trajectories to be exploited. For example a spacecraft with an upgraded NERVA/Pewee-class NTP travelling along an Earth-Jupiter-1I trajectory, would reach 1I/’Oumuamua within 14 years of a launch in 2031. The payload mass to 1I/’Oumuamua would be around 2.5metric tonnes, but even larger masses and shorter mission durations can be achieved with some of the more advanced NTP options studied. In all 4 different proposed NTP systems and 5 different trajectory scenarios are examined.
2.2 NTR Propulsion Options
Four NTR options are considered here and are provided in Table 1.
Table 1 : NTR options with their performance values NTR motor Description Ref. Mass
Upgraded version of the Pewee studied
in the NASA NTR NERVA program
‘50s to early ‘70s
 3250 906 111.7 SNRE Based on the Small Nuclear Rocket
Engine, studied in the Rover program
 2400 900 73 SLHC Square Lattice Honeycomb  2500 970 147.5 SNTP Particle Bed Nuclear Thermal Rocket  800 950 196
It is assumed the s/c is transported to a LEO of 406km by a NASA Space Launch System (SLS) Block 2. It is currently envisaged that an SLS offers a 130metric tonne capability to LEO. We further assume that only one NTR motor is used, and LH2 propellant can be stored for significant durations with no-leakage and with a zero boil-off cryocooler , and further that there are 2 Staged LH2 tanks with an optimum mass ratio. Payload here is understood to mean the total mass of spacecraft after the engines and spent LH2 tanks have been jettisoned. This therefore represents the useful available spacecraft mass. Note howeve\a\r that this mass includes that of any heat shield which may be necessary if a Solar Oberth is involved.
We get Figure (1) for payload mass against ΔV budget.The ratio of dry stage mass to wet stage mass for both stages is assumed to be p=0.035 which was calculated using the same spacecraft/LH2 mass budgets provided in and incorporates the tank insulation and cryocooler mass.
To generate Figure (2) for Ammonia (NH3) propellant a factor of 0.63 is applied to the specific impulses provided for LH2 above. The value of p for LH2 is retained. This is probably an overestimate as the requirements on storage/insulation would be less severe than for LH2, so this Figure (2) is probably a conservative estimate.
The data from Figures (1) & (2) will be used in Section (3.2) as input to generate flight times as a function of payload mass to 1I/‘Oumuamua.
2.3 Trajectory Scenarios
There are five scenarios considered here, which partly correspond to scenarios which have already been presented in the previous literature, but using chemical propulsion:
- Direct from Earth to 1I/’Oumuamua[28-30]
- From Earth to 1I using a Solar Oberth
- From Earth to 1I via Jupiter and Solar Oberth[29-31]
- From Earth to Jupiter to 1I using LH2
- Identical to (4) but using NH3
Figures (3), (4), (5) show examples of trajectory scenarios (2), (3), (4/5) respecitvely for illustration. Note that scenarios(1) & (2) have optimum trajectories every one Earth year, due to the position of the Earth with respect to 1I/‘Oumuamua.Scenarios (3), (4) & (5) have optimum trajectories approximately every Jupiter year, so around 12 Earth years, due to the alignment of Jupiter with ‘Oumuamua. Scenario (3) has optima in 2033, 2045, 2057 and so on. Scenarios (4) & (5) have launch optima in 2031, 2043, 2055 etc.
3.1 Mission Flight Duration and ΔV Budgets for Different Trajectory Scenarios
The available ΔV for NTR is generally larger than that for chemical propulsion.
Figure (6) shows minimum flight duration againt launch date (based on yearly optima) for the years 2025 to 2045 and taking a direct trajectory (scenario (1) ). Four differentΔV budgets are allocated, 25km/s, 30km/s, 35km/s and 40km/s. The s/c is assumed to have already been placed in LEO of altitude 406km. Observe that for ΔV budgets of 25km/s and with launch years > 2035, direct trajectories from LEO to 1I have prohibitively long mission durations.
Minimum flight duration plots against Solar Oberth perihelion solar radial distance are shown for three different ΔV budgets for scenario (2) in Figure (7) and four ΔV budgets for scenario (3) in Figure (8). Scenarios (4) & (5) minimum flight duration against ΔV is shown in Figure (9). For Figures (6), (7), (8) & (9) it is assumed the s/c starts in an LEO of altitude 406km.
Even with the powerful Space Launch System (SLS) Block 2, it can be shown that for scenarios (1), (2) & (4/5), chemical propulsion cannot deliver ΔV’s of the magnitude needed for sensible flight durations but scenario (3) is achievable.
3.2 Payload Masses to ‘Oumuamua Achievable Using NTR
For direct trajectories, i.e. scenario (1), data from Figure (6) can be combined with Figure (1) to give Figure (10).
Data from Figures (7), (8) and (9) can be combined with Figure (1) to give Figures (11)-(13). Figure (11) is scenario (2) and note that the NERVA Pewee-class engine is excluded from this plot because it cannot achieve payload masses > 0kg since the required ΔV budgets are too high.
The results are summarised in Table 2. Regarding these Figures (10)-(13), it can be observed that there is a clear relationship between payload mass and flight time. Hence as one might expect, as flight time goes up so the ΔV reduces enabling higher payload masses (ref. Figure (1)). Generally of the NTR options, the NERVA Pewee-class NTR has the worst performance (in terms of longer flight times and lower payload masses) whereas the Particle Bed (SNTP) has the best performance. As also might be predicted, for those trajectories which employ a Solar Oberth Maneuver, the closer the Solar Oberth to the sun, the better the overall mission capability, though naturally the solar flux and consequent heat shield mass requirement increase.
Scenario (2) enables lower flight durations than scenario (1) but has lower payload mass capability.
Generally scenario (3) enables higher payload masses to 1I than scenarios (1) & (2) but it can also be seen that scenarios (1) & (2) trajectories have yearly optima as opposed to the 12-yearly optima for scenario (3). Furthermore, flight durations are lower for scenarios (1) & (2). Scenario (4) with LH2 tanks is shown in Figure (13) and offers better performance than scenario (3), but without a hazardous close approach to the sun. Scenario (5) with NH3 tanks is shown in Figure (14) and has lower payload masses compared to LH2 as would be expected.
Comparing Figure (13) (scenario (4) using a powered Jupiter GA with LH2 tanks) against Figure (10) (scenario (1), direct transfer), it appears that the former provides significantly higher payload masses, but this is only because the former has a generally lower ΔV mission profile. In fact for equivalent ΔV budgets, these two scenarios have the same payload mass but scenario (4) gives a significantly lower flight duration. So for instance, with ΔV= 25km/s, scenario (1) gives a duration of 37 years whereas scenario (4) with LH2 has a duration of 14 years. However there are two key disadvantages with scenario (4) and they are first that it involves a journey to Jupiter which requires the LH2 to be stored without significant leakage and with zero boil-off (so with a cryocooler) and second the launch optima are at twelve year intervals, between which trajectories are not viable.
Table 2 Results Summary Trajectory
1 Direct from
LEO to 1I
Every year 2.9
2.7 11.1 38.0 26.2 17.8 2 LEO to Solar
Oberth to 1I
Every year 0 2.9
1.3 6.5 25.3 21 19.4 3 LEO to
7.8 15.2 72.8 32 43.1 4 LEO to
Jupiter to 1I
8.6 10.7 29.0 17.6 39.7 5 LEO to
Jupiter to 1I
3.6 13.8 29.0 23.6 39.7
In this paper, we investigated the use of NTR for chasing interstellar objects, once they have left the inner solar system. We used the example of 1I/’Oumuamua for illustration.
We identified several advantages of using NTR for a mission to ‘Oumuamua and similar interstellar objects. First, due to the higher Isp, flight durations can be considerably reduced to < 15 years compared to > 20 years for chemical propulsion. The higher Isp also implies that generally the payload masses to 1I/’Oumaumua are considerably larger than for chemical propulsion. This translates to payload masses of 1000’s of kg as opposed to 100’s of kg for chemical.
In terms of the trajectories, direct trajectories are possible, which significantly reduce the complexity of missions. Direct trajectories also mean that the optimum launch windows arrive once a year, when the Earth and ‘Oumuamua are appropriately aligned. Second, in contrast to chemical propulsion, the arrival velocities are much lower approx. 18km/s, compared to 30km/s with chemical. Lower velocities allow for longer observation times during the encounter and thereby a higher science return. Trajectories without a Solar Oberth maneuver also have the advantage of a lower degree of uncertainty. One of the caveats of the Solar Oberth maneuver is that the errors or uncertainties in the burn at Perihelion have a disproportionate influence on the solar system escape trajectory. Also, the heat shield is not required, thereby saving mass. If a Solar Oberth is utilized, they can be farther from the sun (perihelia for the SO can be > 10Solar Radii) compared to chemical propulsion (where perihelia for the SO are typically< 10 Solar Radii), which reduce the requirements for the heat shield, as the solar irradiation per area diminishes with 1/r².
A major drawback of the NTR trajectories which employ a trip to Jupiter, is the high relative velocity of the s/c as it approaches 1I/’Oumuamua (from Table 2 around 40km/s).
However, though not studied, there is the potential for slowing down as the target is approached and even to perform a rendezvous, though this would be after a long flight duration and so contingent on nearly zero-leakage and zero boil-off LH2 tanks and cryocoolers.
The direct trajectory option (scenario (1)) would not require long LH2 storage durations and for reasons elucidated above is possibly one of the preferred options. However if mission duration is important, scenario (4), Earth-Jupiter-1I, is possibly the preferred choice.
One limitation of our study is that the payload is considered as a black box, and potential constraints from the spacecraft with its instrumentation are not taken into consideration. Such constraints might be related to compatibility issues between NTR and certain spacecraft instruments.
To summarize, our findings indicate that NTR for missions to interstellar objects would have a significant effect on the duration of a mission to such objects (trip times can be even cut in half) and allow for payload masses of an order of magnitude higher than for chemical propulsion. Hence, NTR would be a game changer for missions chasing interstellar objects, when they are on their way out of the solar system. Future work should explore the use of NTR for rendezvous and even sample return missions, which are feasible with NTR.
In this paper, we examined the use of Nuclear Thermal Propulsion (NTP) for missions to interstellar objects, exemplified by 1I/’Oumuamua. Four different proposed NTP options are analysed, ranging from NERVA-based designs to more advanced NTP. Using the OITS trajectory optimization tool, we find that NTP would allow for simpler and more direct, time-saving trajectories to 1I/’Oumuamua. Significant savings in terms of mission duration (14 years for a launch in 2031) are identified. Payload masses on the order of 1000s of kg, compared to 100s of kg using a Space Launch System launcher would be feasible. We conclude that NTP would be a game changer for chasing interstellar objects on their way out of the solar system, drastically reducing trip times and increasing payload masses. Future work should explore rendezvous missions using NTP as well as the feasibility of using NTP for reactive missions, where interstellar objects are discovered early.
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 Power||0.198-0.065 GW|
|Exhaust velocity||See Table|
|Engine mass||10 tonne|
|Exhaust Velocity||8,093 m/s|
|Specific Impulse||825 s|
|Thrust Power||0.2 GW|
|Mass Flow||6 kg/s|
|Total Engine Mass||10,000 kg|
|Remass Accel||Thermal Accel:|
|Specific Power||50 kg/MW|
|Resuable Nuclear Shuttle [+]|
|Exhaust Velocity||8,000 m/s|
|Specific Impulse||815 s|
|Thrust Power||1.4 GW|
|Mass Flow||43 kg/s|
|Wet Mass||170,000 kg|
|Dry Mass||30,000 kg|
|Mass Ratio||5.67 m/s|
|Widmer Mars Mission [+]|
|Exhaust Velocity||8,000 m/s|
|Specific Impulse||815 s|
|Thrust Power||2.3 GW|
|Mass Flow||72 kg/s|
|Wet Mass||400,000 kg|
|Dry Mass||150,000 kg|
|Mass Ratio||2.67 m/s|
|HELIOS 2nd Stage [+]|
|Propulsion System||NTR Solid|
|Exhaust Velocity||7,800 m/s|
|Specific Impulse||795 s|
|Thrust Power||3.8 GW|
|Mass Flow||126 kg/s|
|Wet Mass||100,000 kg|
|Dry Mass||6,800 kg|
|Mass Ratio||14.71 m/s|
|Atomic V-2 [+]|
|Propulsion System||NTR Solid|
|Exhaust Velocity||8,980 m/s|
|Specific Impulse||915 s|
|Thrust Power||4.7 GW|
|Mass Flow||117 kg/s|
|Total Engine Mass||4,200 kg|
|Wet Mass||42,000 kg|
|Dry Mass||17,000 kg|
|Mass Ratio||2.47 m/s|
|Specific Power||1 kg/MW|
|Exhaust Velocity||9,200 m/s|
|Specific Impulse||940 s|
|Thrust Power||512 MWt|
|Mass Flow||12.5 kg/s|
|Total Engine Mass||3,240 kg|
|Fissle Loading||0.25 g U per cm3|
|Max Fuel Temp||2940 K|
|U-235 Mass||36.8 kg|
|Chamber Pressure||1000 psi|
|Remass Accel||Thermal Accel:|
(inc. skirt ext)
|Nozzle Exit Dia||1.87 m|
|Specific Power||6.3 kg/MW|
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 klbf) "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 klbf) 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 klbf) 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.
|Exhaust Velocity||8,770 m/s|
|Specific Impulse||894 s|
|Thrust Power||145 MWt|
|Mass Flow||3.8 kg/s|
|Total Engine Mass||1,770 kg|
|Max Enrichment||93% U-235 wt|
|Num Fuel Elements||260|
|Fissle Loading||0.6 g U per cm3|
|Max Fuel Temp||2736 K|
|U-235 Mass||27.5 kg|
|Remass Accel||Thermal Accel:|
|Specific Power||12.2 kg/MW|
|Engine Length||6.19 m|
|Exhaust Velocity||8,829 m/s|
|Specific Impulse||900 s|
|Thrust Power||367 MWt|
|Mass Flow||8.4 kg/s|
|Total Engine Mass||2,400 kg|
|Max Enrichment||93% U-235 wt|
|Num Fuel Elements||564|
|Fissle Loading||0.6 g U per cm3|
|Max Fuel Temp||2,726 K|
|U-235 Mass||59.6 kg|
|Remass Accel||Thermal Accel:|
|Specific Power||6.5 kg/MW|
|Engine Length||4.46 m|
|Fuel Length||0.89 m|
|Exhaust Velocity||9,120 m/s|
|Specific Impulse||930 s|
|Thrust Power||2.0 GW|
|Mass Flow||49 kg/s|
|Total Engine Mass||9,000 kg|
|Remass Accel||Thermal Accel:|
|Specific Power||4 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 Exhaust Velocity 9,810 m/s Specific Impulse 1,000 s Thrust 134,400 N Thrust Power 0.7 GW Mass Flow 14 kg/s Total Engine Mass 32,546 kg T/W 0.42 Frozen Flow eff. 73% Thermal eff. 96% Total eff. 70% Fuel Fission:
Reactor Solid Core Remass Liquid Hydrogen Remass Accel Thermal Accel:
Thrust Director Nozzle Specific Power 49 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.
This is from DESIGN COMPARISON OF NUCLEAR THERMAL ROCKET CONCEPTS (2019)
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:
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).
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.
- Be: beryllium
- BeO: beryllium oxide
- B4C: boron carbide
- Graph.: graphite
- Hf: hafnium
- Hyd: liquid hydrogen
- Inc718: inconel 718
- Mo: molybdenum
- SS-316: marine grade stainless steel
- W, W10: tungsten
- W5Re, W25Re: tungsten-rhenium alloy
- ZrC, ZrC40: zirconium carbide
- ZrH1.6, ZrH2: zirconium hydride
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.
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.
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.
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.
|Thrust Power||14.0-4.6 GW|
|Exhaust velocity||See Table|
|Engine mass||5 tonne|
|Exhaust Velocity||8,093 m/s|
|Specific Impulse||825 s|
|Thrust Power||14.2 GW|
|Mass Flow||432 kg/s|
|Total Engine Mass||5,000 kg|
|Remass Accel||Thermal Accel:|
|Dumbo Model A|
|Engine mass||0.7 tonne|
|Propellant mass flow||52 kg/sec|
|Exhaust velocity||7,700 m/sec|
|Engine Height||0.6 m|
|Engine Radius||0.3 m|
|Engine Volume||0.2 m3|
|Dumbo Model B|
|Engine mass||2.8 tonne|
|Propellant mass flow||460 kg/sec|
|Exhaust velocity||7,700 m/sec|
|Engine Height||0.6 m|
|Engine Radius||1.0 m|
|Engine Volume||1.8 m3|
|Dumbo Model C|
|Engine mass||2.1 tonne|
|Propellant mass flow||48 kg/sec|
|Exhaust velocity||8,300 m/sec|
|Engine Height||0.6 m|
|Engine Radius||0.4 m|
|Engine Volume||0.3 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.
|Exhaust Velocity||9,530 m/s|
|Specific Impulse||971 s|
|Thrust Power||1.6 GW|
|Mass Flow||35 kg/s|
|Total Engine Mass||1,700 kg|
|Remass Accel||Thermal Accel:|
|Specific Power||1 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 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.
|Engine Mass||6,803 kg||1,500 kg|
|Thrust (Vac)||333.6 kN||392.8 kN|
|Specific Impulse||850 s||1,000 s|
|Burn Time||1,200 s||449 s|
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.
|Diameter||4.25 m||2.03 m||8.70 m|
|Thrust (Vac)||392.8 kN||735.5 kN||2,451.6 kN|
|Specific Impulse||1,000 s||1,000 s||1,000 s|
|Engine Mass||1,500 kg||2,500 kg||8,300 kg|
|Burn Time||449 s||357 s||493 s|
One of the design goals was to make the engine work with low enriched uranium (LEU) instead of weapons-grade uranium (HEU) like older designs. LEU is harder to make work in a nuclear thermal engine, but weapons-grade uranium in civilian hands make the military very nervous.
INsTAR Fuel LEU
Engine Power 1.76 GW Mass flow rate 25.15 kg/s Exhaust Vel 10,610 m/s Isp 1,082 sec T/W 9.5 Thrust 270,000 N Engine Mass ~2,900 kg
The development of a low enriched, nuclear thermal rocket (NTR) has become a necessity for more effective deep space travel. An NTR provides a significantly higher specific impulse (Isp) than chemical rockets. Starting in the 1960s the United States began the Nuclear Engine for Rocket Vehicle Applications (NERVA) program. The overall design of the NTR described in this paper, INsTAR, incorporates many elements from the NERVA prototypes. While many NERVA engines incorporated a high enriched uranium fuel, the low enriched uranium fuel concepts have proven to be feasible. The overall design objective was to develop a full core design for an NTR that addresses the problems found in the Space Nuclear Thermal Propulsion Project core and improves upon the Isp and thrust to weight ratio observed during the project. The core design is characterized with respect to neutronics, thermal hydraulics, and propulsive performance. A critical component of the design was the study of materials utilized in the core and their compatibilities at high operating temperatures.
Nuclear thermal propulsion (NTP) is a leading, inspace design to reduce travel time for astronauts and carry larger payloads than contemporary chemical rockets. Efforts to develop nuclear thermal rockets (NTRs) began in the 1960s. One such effort, the Nuclear Engine for Rocket Vehicle Applications (NERVA), was a nuclear thermal rocket engine development program that ended in 1972 (Ref. 1). An example of a rocket engine developed by the NERVA program is seen in Fig. 1. NERVA ultimately demonstrated that NTRs were a feasible technology for space exploration. One of the engine designs during this time met all requirements for a mission to Mars. Since then, it has been determined that NTRs are essential for regular trips to Mars and human exploration of further planets. Approximately twenty experiments of different types of nuclear reactors for space applications were completed.1
Since the feasibility of low enriched uranium (LEU) for NTR applications has been demonstrated, two NERVA-inspired LEU concepts have been under investigation. These concepts are the Space Capable Cryogenic Thermal Engine (SCCTE) and the Superb Use of Low Enriched Uranium (SULEU). Both designs are based on the same geometry as the NERVA engine and use similar materials and components. The difference between these two concepts are in the fuel design. The SCCTE fuel design utilizes enriched Tungsten-184 ceramic-metallic fuel. The SULEU fuel design utilizes (U,Zr)C-Graphite composite fuel that has been the typical fuel design of previous NTRs. Both concepts utilize LEU fuel with an enrichment of 19.75% U-235. They also are designed to operate with a specific impulse of 900 seconds at a thrust of 35k lbf.3
Particle bed reactor (PBR) concepts are used in space applications and utilize grain sized spherical particles composed of fuel and cladding. Thousands of fuel particles are arranged in an array and can be in many different configurations based on desired properties. Several layers of fuel elements create the reactor and gas, typically hydrogen, radially flows through the fuel elements to cool the system. The hydrogen is heated and expands as it is ejected axially. The fuel's very high surface-to-volume ratio is desirable as this allows for high heat transfer efficiencies and large power densities.1 The PBR's compact design and high power density are ideal attributes for space applications.
The coolant and propellant used in the PBR design is liquid hydrogen. Liquid hydrogen has the lowest molecular weight of all substances, leading to a higher specific impulse than other propellants. It is cryogenic and therefore should be handled with extreme care and stored below 423 °F. Shielding and insulation are required to keep the propellant from evaporating or boiling. The shielding will protect the propellant from the radiant heat of the sun and the insulation will provide protection from other sources of heat like air friction during the flight through the atmosphere.4 Liquid hydrogen is the signature fuel used by the American space program and works well with the proposed NTR design, INsTAR.
Specific impulse (Isp) and thrust to weight ratios (T/W) are important performance metrics for space exploration. The NTR is able to achieve an Isp exceeding twice that of a traditional chemical rocket, helping account for the lower T/W than chemical engines.1 The proposed design improves upon the NERVA derived Space Nuclear Thermal Propulsion (SNTP) engine. The earliest NTR designs achieved a T/W of 4:1 to 8:1 with SNTP achieving a T/W of 16.3:1. The SNTP design incorporates a PBR.1 The fuel elements were composed of high enriched uranium (HEU) with a tungsten (W) cladding that was primarily W-184. The W-184 cladding allows the fuel to operate at much higher temperatures while still providing a lower thermal neutron absorption cross-section.5
The SNTP engine was selected as the design basis for the LEU NTR, INsTAR, for its high T/W and its Isp of 930 seconds. Despite this success, the engine endured failures that are addressed in the final design of INsTAR. Initial nuclear element test runs yielded power irregularities caused by fuel element failure. The fuel had migrated within the fuel bed, creating regions of maximum stress, not accounted for in previous thermal calculations. The hot frit had split, and the cold frit had warped due to stresses from fuel bed thermal expansion. Later particle bed nuclear tests had derived the melting temperature of the UC2 to be 2500K, below that of the operating temperatures of the LEU NTR.1
The overall goal of this project is to improve on the issues that were experienced during the SNTP program and to expand on research that was done previously. Some of these issues included fuel migration and the inability to achieve the high temperatures desired.1 The fuel design used was chosen to prevent the issues seen in the SNTP engine and to increase thermal capabilities. The end goals of this project are to achieve an Isp of approximately 1000 sec. and a T/W between 8:1 and 10:1 with a pressure drop across the core that is less than 1000 psi. MCNP 6.1 (Ref. 6) and thermal hydraulics calculations were used to validate the core design. Compatibility of the materials at high temperatures was investigated. The rocket performance was characterized through Isp and T/W calculations.
III. REACTOR CONFIGURATION
The proposed reactor is an improvement on the PBR design that emerged from the SNTP project. The system design is outlined in the following sections.
III.A. Fuel Pellet
The fuel pellets consist of spherical particles with an internal kernel of uranium (U) metal enriched to 12.5% U-235. The kernel of U metal is clad in a 0.2 mm thick tungsten-rhenium alloy for a total outer diameter of 2 mm, as seen in Fig. 2. There is a 0.1 mm gap between the fuel and the cladding to accommodate for fission gases and to account for thermal expansion of the fuel as it is heated and then melts.
The cladding is composed of 90% W-184 with 5% natural W and 5% rhenium (Re). Adding 5% Re to the cladding improves ductility and strength of the overall material and will improve the ability to spherically encase the fuel.7 Tungsten alone isn't ductile enough to successfully encase uranium during operation of INsTAR.
The fuel particle design allows the U metal to operate in a molten state while the cladding is still in a stable solid state. This is because the melting temperature of U is 1135 °C (Ref. 8) whereas the melting temperature of W is 3422 °C (Ref. 9) and the melting temperature of Re is 3185 °C (Ref. 10). From a phase diagram, the melting temperature of the WRe alloy is about 3300 °C (Ref. 11). The material interaction we would potentially see between the cladding and fuel would be a recrystallization of the two materials. This shouldn’t be too much of a concern given that we are only burning for several minutes at a time. The fuel design allows the rocket to safely operate at temperatures up to approximately 3500 K.
III.B. Fuel Element Design
The fuel particles will be arranged in a hexagonal close-packed design, as seen in Fig. 3, with the particles sintered together to form a fuel element which will prevent particle migration. The particles will form a cylinder with a conical shaped channel in the center to allow for the hydrogen gas to flow out towards the nozzle, as seen in Fig. 4. There is a hot frit that lines this channel, composed of porous graphite. The cold frit is also composed of porous graphite and lines the outside of the fuel particles. The fuel particles are surrounded by graphite moderator, as seen in Fig. 4.
III.C. Core Description
The core consists of identical fuel elements surrounded by a beryllium reflector. The center of the fuel element is a conical-shaped channel allowing for flow of the H2 propellant, as seen in Fig. 4. The hot frit is surrounded by a 10 cm wide conical particle bed tapering to 5 cm consisting of the sintered fuel pellets described in the fuel pellet section. These particles are located between the cold frit and hot frit in Fig. 4. The particle bed is surrounded by a 15 cm wide hexagonal shaped graphite moderator. Fuel elements are arranged to fit 37 identical fuel elements inside of a beryllium (Be) reflector as shown in Fig. 5. Six beryllium control drums were added to the design to account for the excess reactivity. One third of each control drum is lined with boron carbide, which has a high neutron absorption cross section. The final dimensions of the core were 150 cm tall with a 140 cm diameter. The mass of the core is estimated to be approximately 1500 kg, with 310 kg of U metal. Crosssections of the core can be seen in Fig. 5 and Fig. 6.
IV. CORE CHARACTERIZATION
Several aspects of core performance were evaluated before a final design was reached. MCNP 6.1 (Ref. 6) was used to characterize the neutronic capabilities of the core. Thermal hydraulic calculations were completed to determine pressure drops, heat generation, and the maximum core temperature. Propulsion calculations were performed to determine Isp and the T/W. Both the neutronics and thermal hydraulics calculations required a packing efficiency, which is 74% for hexagonally close-packed systems.
Through neutronics calculations, it was determined that the original design had an extreme amount of excess reactivity. This resulted in the decision to reduce the enrichment of the uranium, making this design more favorable regarding uranium enrichment. The final design had a keff of 1.02134 ± 0.0006. Six beryllium control drums were added to the design for reactivity control. When the 120° arc of boron carbide is turned towards the core, the reactivity is reduced, and the reactor becomes subcritical. Fig. 7 shows the effect of the control drums on the reactivity.
IV.B. Thermal Hydraulics
Thermal hydraulics calculations were used to ensure the feasibility of the system and were correlated with the neutronics calculations to ensure the design was consistent. The results of the thermal hydraulic analysis regarding temperature and pressure are summarized in Table 1 below. The maximum temperature within the system is 3301 K. The system requires a mass flow rate of 25.15 kg/s, resulting in a pressure drop across the core of 623 psi, which meets the design criteria. The total power was calculated to be 1.76 GW.
Core inlet and outlet parameters
Parameter Core Inlet Core Outlet Temperature (K) 33 3300 Pressure (psi) 1623 1000
IV. C. Rocket Performance
When a final design was decided upon, the Isp and T/W ratio was calculated, along with the Mach number at several sections of the nozzle. A schematic of a nozzle can be seen in Fig. 8. The hydrogen exits the core and mixes in a conical section between the core and nozzle which results in a homogeneous hydrogen temperature. From there, the gas enters the nozzle with a Mach number less than one and reaches a Mach number of one at the throat of the nozzle, as seen in Fig. 8. From here, the gas continues to speed up and exits the nozzle at a supersonic velocity and a Mach number greater than one. The gas exits the nozzle with a velocity of 10,610 m/s and a Mach number of 5.66. The Isp for this design is 1,082 seconds and the T/W is 9.5 for a thrust of 60,000 lbf. The Isp calculation does not account for the hydrogen dissociation that would occur at the operating temperature of this design. Accounting for hydrogen dissociation will increase the Isp due to a lowered average molecular weight. The Isp determined for this design is higher than several other design concepts. This is due to the fact that a higher maximum operating temperature is allowed because the tungsten-rhenium cladding has a high melting temperature. For comparison, the SNTP Isp was 930s, and the NERVA Pewee had an Isp of 901s.1
This analysis has demonstrated that the preliminary design of the INsTAR rocket is feasible. Further optimization and design changes need to be implemented to increase the usefulness of the design, such as increasing the T/W. The Isp, T/W, and pressure drop across the core met design criteria and further support the feasibility of this design. An Isp of 1,082 seconds is higher than any other previous NTR designs. This high specific impulse is largely due to the innovative use of a tungsten-rhenium cladding, which allows the fuel to operate in a molten state and the propellant to be heated to a higher temperature than is otherwise possible. The T/W is larger than the values achieved by similar sized NTRs. For future work, determination of the hydrogen dissociation rate will allow for more accurate Isp calculations. At this point in time, from a material stand point, the design is compatible from all angles. There needs to be optimization for materials and the interactions that they undergo. This will come from testing and expanding on any new research that is available. Major design components that need to be addressed further include the cladding stability at very high temperatures and the flow of high velocity gases through particle fuel beds. In addition, the MCNP models have fuel elements which approximate the particle bed fuel as a porous medium. Though a design such as this one has not been fully operable, there are many things to learn and overcome. By addressing these minor issues, it will be possible to make a low enriched nuclear thermal rocket for deep space exploration. Our research into this NTR has proven that LEU nuclear thermal propulsion is possible with further optimization and testing. As a future goal, the hope is that we will eventually have a working rocket core to properly and safely achieve propulsion, expanding opportunities for deep space missions.
- R. A. HASLETT, “Space Nuclear Thermal Propulsion Program Final Report,” Grumman Aerospace Corporation (1995).
- NASA ON THE COMMONS, Flickr (2018).
- P. VENNERI and Y. KIM, “Advancements in the Development of Low Enriched Uranium Nuclear Thermal Rockets,” Energy Procedia, 131 (2017).
- B. DUNBAR, NASA – Liquid Hyrogen – the Fuel of Choice for Space Exploration, NASA.gov (2018).
- R. L. MACKLIN, D. M. DRAKE and E. D. ARTHUR, “Neutron-Capture Cross Sections of the Tungsten Isotopes 182W, 183W, 184W, and 186W from 2.6 to 2000 keV,” Oak Ridge National Laboratory (1982).
- T. GOORLEY, et.al., “Initial MCNP6 Release Overview,” Nuclear Technology, 180, p. 298-315 (2012).
- M. S. EL-GENK, “A review of refractory metal alloys and mechanically alloyed-oxide dispersion strengthened steels for space nuclear power systems,” Science Direct (2004).
- Uranium – Element information, properties and uses | Periodic Table, (n.d.).
- Tungsten – Element information, properties and uses | Periodic Table, (n.d.).
- Rhenium – Element information, properties and uses | Periodic Table, (n.d.).
- U. RAVI KIRAN, et.al., “Refractory metal alloying: A new method for improving mechanical properties of tungsten heavy alloys,” Journal of Alloys and Compounds, 709 (2017).
- C. NEWEY and G. WEAVER, Materials Principles and Practice, M. KEYNES, Open University, Materials Department (1991).
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. But see SSailor67's comment below.
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.
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
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!
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.
|Thrust (vac) (kN)||35.28||68|
|Propellant||H2 + Hexane||H2|
|Propellant Mass Flow (kg/s)||~4||~7.1|
|Specific Impulse (vac) (s)||~900||~920|
|Core outlet temparture (K)||3,000||2,800 to 2,900|
|Chamber Pressure (105 Pa)||70||60|
|U235 enrichment (%)||90||90|
|Fuel Element Form||Twisted ribbon||Twisted ribbon|
|Generated electrical power (kW)||N/A||50|
|Working fluid for power loop|
(% by mass)
|N/A||93% Xe + 7% He|
|Max temp for power loop (K)||N/A||1,500|
|Max press for power loop (105 Pa)||N/A||9|
|Working fluid flow rate (kg/s)||N/A||1.2|
|Thermal power - propulsion mode (MW)||196||340|
|Thermal power - power mode (MW)||N/A||0.098|
|Core length (mm)||800||700|
|Core diameter (mm)||500||515|
|Engine length (mm)||3,700||No Data|
|Engine diameter (mm)||1,200||No Data|
|Lifetime - propulsion mode (h)||1||5|
|Lifetime - power mode (yr)||N/A||2|
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 power 1,650 MW Exhaust velocity 9,420 m/s Specific impulse 960 s Thrust 330,000 N Engine mass 5,260 kg T/W 6.4
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.
|Engine Mass||835 kg|
|Full Thrust||49,000 newtons|
|Full Isp||1,210 sec|
|Single-H Thrust||9,800 newtons|
|Single-H Isp||1,350 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.
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 Velocity||9,221 m/s|
|Specific Impulse||940 s|
|Thrust Power||0.3 GW|
|Mass Flow||7 kg/s|
|LANTR LOX mode|
|Exhaust Velocity||6,347 m/s|
|Specific Impulse||647 s|
|Thrust Power||0.6 GW|
|Mass Flow||29 kg/s|
|Remass||Hydrogen + Oxygen|
|Remass Accel||Thermal Accel:|
|Nuclear DC-X NERVA|
|Exhaust Velocity||9,810 m/s|
|Specific Impulse||1,000 s|
|Thrust Power||27.3 GW|
|Mass Flow||567 kg/s|
|Specific Power||7 kg/MW|
|Exhaust Velocity||5,900 m/s|
|Specific Impulse||601 s|
|Thrust Power||49.2 GW|
|Mass Flow||2,827 kg/s|
|Remass Accel||Thermal Accel:|
|Specific Power||4 kg/MW|
|Total Engine Mass||199,600 kg|
|Wet Mass||460,000 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
|Bimodal NTR Solid (NASA)|
|Propulsion System||NTR Solid Bimodal|
|Exhaust Velocity||8,980 m/s|
|Specific Impulse||915 s|
|Thrust Power||0.9 GW|
|Mass Flow||22 kg/s|
|Total Engine Mass||6,672 kg|
|Remass Accel||Thermal Accel:|
|Wet Mass||80,000 kg|
|Dry Mass||26,830 kg|
|Mass Ratio||2.98 m/s|
|Specific Power||7 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.
Dual-mode Fission Exhaust Velocity 9,810 m/s Specific Impulse 1,000 s Thrust 124,700 N Thrust Power 0.6 GW Mass Flow 13 kg/s Total Engine Mass 33,000 kg T/W 0.39 Thermal eff. 94% Total eff. 94% Fuel Fission:
Reactor Solid Core Remass Liquid Hydrogen Remass Accel Thermal Accel:
Thrust Director Nozzle Specific Power 54 kg/MW Special Bimodal Thermal Electrical eff. 19% Electrical Power 60 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).
Pebble-bed Fission Reactor Exhaust Velocity 9,810 m/s Specific Impulse 1,000 s Thrust 172,700 N Thrust Power 0.8 GW Mass Flow 18 kg/s Total Engine Mass 58,000 kg T/W 0.30 Thermal eff. 94% Total eff. 94% Fuel Fission:
Reactor Solid Core Remass Liquid Hydrogen Remass Accel Thermal Accel:
Thrust Director Nozzle Specific Power 68 kg/MW Special Bimodal Electrical Power 60 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%).
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.
|Exhaust Velocity||9,810 m/s|
|Specific Impulse||1,000 s|
|Thrust Power||68.7 MW|
|Mass Flow||1 kg/s|
|Total Engine Mass||200 kg|
|Remass Accel||Thermal Accel:|
|Specific Power||3 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.
|Exhaust Velocity||12,750 m/s|
|Specific Impulse||1,300 s|
|Thrust Power||15.0 MW|
|Mass Flow||0.18 kg/s|
|Total Engine Mass||200 kg|
|Remass Accel||Thermal Accel:|
|Specific Power||13 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 Velocity||17,660 m/s|
|Specific Impulse||1,800 s|
|Thrust Power||15.0 MW|
|Mass Flow||0.10 kg/s|
|Total Engine Mass||10,000 kg|
|Remass Accel||Thermal Accel:|
|Specific Power||666 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 1|
|Exhaust Velocity||16,000 m/s|
|Specific Impulse||1,631 s|
|Thrust Power||56.0 GW|
|Mass Flow||438 kg/s|
|Total Engine Mass||70,000 kg|
|Remass Accel||Thermal Accel:|
|Specific Power||1 kg/MW|
|Liquid Core 2|
|Exhaust velocity||14,700 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 Velocity||19,620 m/s|
|Specific Impulse||2,000 s|
|Thrust Power||0.2 GW|
|Mass Flow||1 kg/s|
|Total Engine Mass||1,000 kg|
|Remass Accel||Thermal Accel:|
|Specific Power||5 kg/MW|
|Exhaust Velocity||10,300 m/s|
|Specific Impulse||1,050 s|
|Thrust Power||56.6 GW|
|Mass Flow||1,068 kg/s|
|Total Engine Mass||9,000 kg|
|Wet Mass||226,000 kg|
|Dry Mass||45,000 kg|
|Mass Ratio||5.02 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
|Reactor inner diameter||1 m|
|Reactor outer diameter||2 m|
|Reactor inner length||3 m|
|Reactor outer length||4 m|
|Engine length||13 m|
(no shadow shield)
(with shadow shield)
(no shadow shield)
(with shadow shield)
|Engine pressure||500 atm|
|Exhaust velocity||19,600 m/s|
|Engine power||1,500 MWth|
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:  help keep the blasted uranium drops from splattering on the walls and  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.
|Thrust Power||1.6 GW|
|Exhaust velocity||9,800 to|
|Propellant mass flow||30 kg/sec|
|Reactor thermal power||1,400 to|
|Total engine mass||6.83 tonne|
|Fuel element mass total||1.35 tonne|
|Forward reflector mass||0.60 tonne|
|Aft reflector mass||0.51 tonne|
|Radial reflector mass||2.47 tonne|
|Radiation shield mass||0.9 tonne|
|Total reactor mass||5.83 tonne|
|Remass Accel||Thermal Accel:|
|Specific Power||4 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).
|Forward||Beryllium oxide||15 cm||0.60 tonne|
|Aft||C-C Composite||25 cm||0.51 tonne|
|Radial||Beryllium oxide||15 cm||2.47 tonne|
|CORE: 2000 fuel elements|
|Fuel channel per element||12 to 32|
|Hydrogen channel per element||12 to 32|
|Critical mass||20 kg|
|Hydrogen pressure||100 atm|
|UF4 pressure||100 atm|
|Fuel center temperature||4,500 K|
|Pump Flowrate (Total)||75.20 lbm/s|
|Pump Discharge Pressure||3,924 psia|
|Turbopump RPM||70,000 RPM|
|Turbopump Power (each)||9,836 HP|
|Turbine Inlet Temperature||481 deg-R|
|Turbine Pressure Ratio||1.69|
|Turbine Flow Rate (each)||33.77 lbm/s|
|Reactor Thermal Power||1,769 MW|
|Fuel Element and Reflector Power||1,716 MW|
|Nozzle Chamber Temperature||5,580 deg-R|
|Chamber Pressure (Nozzle Stagnation)||1,500 psia|
|Nozzle Expansion Area Ratio||500:1|
|Vacuum Specific Impulse (Delivered)||997.8 sec|
|Nozzle-con (total)||30.05 MW|
|Nozzle-div (total)||22.97 MW|
|Reflector (total)||35.0 MW|
|Typical NVTR Engine Parameters|
|Nozzle Area Ratio||500|
|Fuel Pressure||100 atm|
|Average Fuel Temperature||4000 K|
|Maximum Element Heat Flux||420 W/cm2|
|Nomial Element Length||150 cm|
|Fuel Volume Fraction||0.15|
|Coolant Volume Fraction||0.15|
|Moderator Volume Fraction||0.70|
|Fuel Element Power||0.9 MWt|
|Element Heat Transfer Area||2141 cm2|
|Reactor Core L/D||1.5|
|Fuel Channel Diameter||0.142 cm|
|Fuel Channel Sectional Area||0.0158 cm2|
|Total Fuel Channel Area Per Element||0.505 cm2|
|Fuel Element Sectional Area||3.464 cm2|
|Element Diameter (across flats)||2.2 cm|
|Coolant Channel Diameter||0.142 cm|
|Coolant Channel Sectional Area||0.0158 cm2|
|Total Coolant Channel Area Per Element||0.505 cm2|
|Core Volume||1.2 m3|
|Core Volume Density||1,500 MW/m3|
|Fuel Element Mass, Total||1.35 MT|
|Forward Reflector Mass||0.60 MT|
|Aft Reflector Mass||0.51 MT|
|Radial Reflector Mass||2.47 MT|
|Radiation Shield Mass||0.90 MT|
|Total Reactor Mass||5.83 MT|
|Misc. Engine Components Mass||0.9 MT|
|Total Engine Mass||6.83 MT|
|Fuel||Uranium 235 + Zirconium Carbide|
|Low Specific Impulse|
|Exhaust Velocity||11,300 m/s|
|Specific Impulse||1,150 sec|
|Uranium Fuel Loss||20 kg/min|
|Engine Pressure||76 atm|
|High Specific Impulse|
|Exhaust Velocity||11,800 m/s|
|Specific Impulse||1,200 sec|
|Uranium Fuel Loss||28 kg/min|
|Engine Pressure||200 atm|
This is from Research on Uranium Plasmas and their Technological Applications (1971). Page 29: The Colloid-Core Concept—A Possible Forerunner for the Gaseous Core.
Gas-Core nuclear-thermal-rocket engines are attractive since they are free from the exhaust velocity limits of solid-core engines. But great galloping galaxies, are those gas-core engines ever an engineering nightmare to design!
So Drs. Stefanko and Dickson of the Westinghouse Electric Corp. Astronuclear Laboratory had the thought it might be worth-while to develop some kind of forerunner engine to the full gas core. An engine whose thrust and specific impulse-exhaust velocity was only mid-way between solid-core and gas-core, but was a heck of a lot easier to design. This will allow spacecraft with nuclear engines superior to solid core NTR to be made now, and give the engineers some breathing room to figure out how to make a real live honest-to-Johnny gas-core engine.
Chemical rocket engines have an exhaust velocity that maxes out at about 4,400 m/s. A good solid-core NTR can double that, about 8,800 m/s. A good open-cycle NTR can manage an exhaust velocity of something like 40,000 m/s without even trying.
Stefanko and Dickson figured out an engine that can manage an exhaust velocity of about 11,800 m/s, with a thrust of around 445,000 Newtons. Which was more than a solid-core, less than a gas-core, and much simpler to engineer.
The researcher found two major difficulties in engineering gas-core NTR: nuclear fuel loss and large engine size. Both of these are because the core is gaseous.
Nuclear Fuel Loss
Since rocket engines need a hole in the bottom to let the exhaust out, the challenge is to allow all of the exhaust to escape but none of the uranium or plutonium. This is tricky since both the uranium fuel and the hydrogen propellant is gaseous, at an incredibly high temperature + pressure, moving at high velocity, and being held in the functional equivalent of an upside-down coffee mug.
You want the propellant to escape because that makes the thrust, which is the entire function of a rocket engine. You do NOT want the uranium fuel to escape because  it turns the exhaust plume into a cloud of glowing radioactive death and  that stuff is expensive, you can't afford to let any un-burnt uranium to get away.
Solid-core NTR prevent the fuel from escaping by making it non-gaseous, at a pathetically low temperature + pressure, not moving at all, and functionally bolted to the upside-down coffee mug so it can't fall out. The price is a lack-luster exhaust velocity.
Closed-cycle gas-core NTR try to confine the gaseous, high temperature + pressure + velocity uranium fuel inside a "lightbulb" to prevent it from fallout out of the upside-down coffee mug. The difficulty is to make a lightbulb (composed of matter) that can withstand the incredibly high temperature and pressure, and allow the heat radiation to leave the lightbulb so it can warm up the propellant. This is trying to have it both ways, since the entire point of a gas-core nuclear rocket was to have the core be gaseous. Having a non-gaseous lightbulb is breaking the rules. In order to prevent the lightbulb from instantly exploding and vaporizing the pressure and the temperature have to be lowered to the point where the exhaust velocity is about half that of an open-cycle gas-core. And even then the engineering challenge of designing an indestructible lightbulb is heart-breaking.
Open-cycle gas-core NTR don't even try to totally confine the uranium, they just try to minimize the loss. They use various clever tricks with vortexes and magnetic fields, but nothing matter-free is going to be 100% successful at confining that radioactive crap.
Large Engine Size
Nuclear fuel will produce about as much energy as a sleeping hippopotamus if you don't have a critical mass. Among other things, you need a large enough amount of uranium inside a small enough volume.
Getting enough uranium inside a small enough volume is relatively easy, if the stuff is solid or liquid. Unfortunately gas is the exact opposite of what we need, it want to get a little uranium inside a volume as possible (this is similar to the way that one small cat can expand itself so it takes up the entire freaking bed). To cram enough gaseous uranium into a given volume requires pressure, lots of it. This raises the engine mass, because pressure vessels whose wall are too thin tend to explode.
Even after you've squeezed the uranium gas as much as possible, you'll find it still lacks the density required to have a critical mass unless the reaction chamber is huge. This increases both the engine's volume size and mass size.
Bottom line is that gaseous uranium has a high critical mass requirement.
So, both of the problems are a direct outcome of the fact that the fuel is gaseous. The researchers asked the question: is to possible to make is not quite gaseous and still get good rocket performance? What if instead of being gaseous, it was instead a colloid?
In this concept, the fuel is not a gas so much as it is a dust-storm. This will allow the fuel to have a density several orders of magnitude (i.e., about a thousand times) as compared to an actual gas.
The nuclear fuel loss problem can be addressed by using the large density difference between the dust particles of uranium and the free atoms of hydrogen.
The large engine size problem is drastically reduced because using fuel particles instead of gaseous fuel allow a critical mass is a much reduced volume. It also dramatically lowers the required pressure, to about half of that required by an open-cycle gas core.
Result: an engine with an exhaust velocity of about 11,800 m/s, with a thrust of around 445,000 Newtons. Which was more than a solid-core, less than a gas-core, and much simpler to engineer.
This concept sounds similar to the pebble-bed reactor engine, but it isn't. The pebble-bed uses a fluidized-bed configuration, and the particles are not completely suspended. This means the pebble-bed fuel is in physical contact with the reactor walls, encouraging them to melt. Which is never a happy occasion.
The concept also sounds similar to the liquid-core reactor engine, but it ain't that either. The liquid core engine uses bubbles of propellant passing through a liquid film of nuclear fuel. This presents flow distribution problems that have been shown to be insurmountable (as of the report date of 1971).
The colloid-core engine uses fluid mechanics to separate the fuel from the propellant in the exhaust stream. The nuclear fuel is suspended in a confined fuel zone (2 in the diagram) to keep it in a critical mass. As mentioned before the pressure required for critical mass was about half that required for a standard open-cycle gas-core, which is good news for the engineers.
According to their mathematical models, the fluid mechanics did a fantastic job of preventing the uranium dust from escaping out the exhaust nozzle. Unfortunately the dust got hot enough to leak uranium vapor from the particle surface, and fluid mechanics didn't do zippidee-doo-dah to stop the vapor from escaping. And the amount of vapor escaping was very significant.
To deal with the uranium vapor problem, they replaced the pure uranium dust fuel with dust composed of a uranium—zirconium carbide compound. A mix of one part uranium carbide to ten parts zirconium carbide (UC2 + 10 ZrC). Zirconium carbide is an extremely hard refractory ceramic material which is highly corrosion resistant and a member of the ultra high temperature ceramics group. Zirconium carbide is already used as a coating for nuclear reactor fuel elements because it has a low neutron absorption cross-section and weak damage sensitivity under irradiation. But the main figure of merit is its low vapor pressure, i.e., it does not vaporize worth a darn.
Mind you, there is still some uranium fuel loss. But it is measured in kilograms per minute, instead of kilograms per second.
Figure 2 above shows the specific impulse given the engine static temperature, the engine pressure in ATMs, and the uranium+zierconium carbide fuel mix. Just ignore all the fuel mixes except for UC2 + 10 ZrC. This figure is mildly interesting.
Figure 3 above shows how much uranium is lost out the exhaust nozzle in kilograms per minute, given engine static temperature, engine pressure in ATMs, and fuel mix. Again ignore all the fuel mixes except for UC2 + 10 ZrC. This figure is mildly interesting.
This is the interesting figure. The curves show specific impulse and uranium fuel loss; given static temperature of engine and engine pressure. The chart assumes UC2 + 10 ZrC fuel, and 445,000 Newtons of thrust (100,000 lbs)
Point "A" (green lines) has specific impulse of 1,150 sec, uranium loss rate of 20 kg/min, static temperature of 3,620°K, and centerline static pressure of 76 atm.
Point "B" (blue lines) has specific impulse of 1,200 sec, uranium loss rate of 28 kg/min, static temperature of 3,950°K, and centerline static pressure of 200 atm.
Table 1 Flow Parameters Sym Parameter Projected Reactor R outer radius of vortext 0.30 m h axial length of chamber 0.30 m ṁ/A total gas flow rate into chamber over A 17.7 kg/m2 sec ρg gas density 1.1 kg/m3 μ gas viscosity 0.04 centripoise ρp material density of particle 8 gm/cm3 D particle diameter 10μ M total powder mass in vortex 30 kg ε void fraction 0.95 r radius 0.13 m ω angular velocity 450 rad/sec
As an example of possible extrapolation from our present experiment to operating conditions of the colloid core reactor, a hypothetical engine will be assessed, with thrust of 20,000 lb force (90,000 Newtons), and specific impulse of 1000 lbf-sec/lbm (Isp 1000 sec, Ve 9,810 m/s). A mass flow rate of 10 kg/sec would then be required. Taking hydrogen as operating medium and assuming a temperature of 3300°K, the dynamic viscosity of the gas flow in the reactor would be about 0.04 centipoise, and gas density at a median cavity pressure of 100 atmospheres would be about 1 kg/m3. A uranium carbide alloy, (1U—lOZr)C (Uranium 235 + Zirconium Carbide), has been proposed as a possible fuel for the colloid core reactor because of its low uranium equilibrium vapor pressure. Material density of this fuel would be about 8 gm/cm3, near the density of zinc. For convenience in relating the behavior of the particulate fuel to our observations of zinc particles, the fuel particle diameter will be taken as 10μ. For this example, the reactor radius is taken as 30 cm, with an axial length of 30 cm, and an inlet velocity of 100 m/sec is given for tangential gas injection.
Critical mass calculations for a cavity radius of 30 cm and axial length of 12 cm indicate a uranium mass requirement of about 3 kg. Critical mass requirement for the longer chamber used in the present example is probably lower than the 3 kg figure, but for a conservative estimate, 3 kg uranium mass, for a total fuel load of 30 kg, will be assumed. Applying the parameters to Eq. (2-4), inner bed radius, angular velocity, and void fraction can be calculated. The various extrapolated flow parameters for the reactor are compared with experimental conditions in Table 1. Note that the particle volume fraction (1 - ε) required by this fuel load is well below the 10% value attained in our present experiments.
The power output required by the engine performance specification is about 600 mw, with a reactor power density of about 104 mw/m3. The average temperature differential between particles and gas is derived in the Appendix, indicating a temperature difference on the order of 20°K between fuel particles and propellant gas for this reactor. The total pressure drop needed to fluidize the particle bed under the prescribed conditions is about 33 atm.
If the zinc particle loss rate given in Fig. 5 is linearly scaled to the larger gas flow rate per unit area under consideration for the reactor, a fuel loss of about 30 gm/sec is predicted, for a uranium loss rate of 3 gm/sec. Vaporization loss for uranium has been estimated at roughly 100 gm/sec at an operating temperature of 3300°K, so that particulate loss will be small in comparison to vaporization.
Our experiments with two component vortex flows, when extrapolated to the operating range of a colloid core reactor, project a reasonable degree of optimism concerning the ultimate feasibility of such a system. The projected reactor size is modest, no problem appears in fluidizing and rotating the required fuel load, and reactor power density is well within the heat-transfer capacity of the fluidized bed. Rough estimates of particulate fuel losses indicate that uranium loss rate will be governed primarily by materials problems, rather than aeromechanical considerations. Many problems concerning the generation and transfer of power within a rotating fluidized bed are yet to be answered, but the basic concept of maintaining a fluidized vortex flow of high-particle density to provide nuclear propulsion appears to be sound.
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.
So crazy engine designers started to look into reactors that were molten under normal operation. Or even gaseous.
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.
|Gaseous Core NTR closed 1|
|Exhaust Velocity||20,405 m/s|
|Specific Impulse||2,080 s|
|Thrust Power||4.5 GW|
|Mass Flow||22 kg/s|
|Total Engine Mass||56,800 kg|
|Remass Accel||Thermal Accel:|
|Specific Power||13 kg/MW|
|Gaseous Core NTR closed 2|
|Thrust Power||0.6 to 231 GW|
|Exhaust velocity||10,800 to 31,400 m/s|
|Thrust||117,700 to 14,700,000 n|
|Engine mass||30 to 300 tonne|
|Engine T/W||0.4 to 5.0|
|Operating Pressure||400 to 1600 atm|
|NASA report nuclear lightbulb|
|Thrust Power||3.7 GW|
|Engine Power||4.6 GW|
|Exhaust velocity||18,300 m/s|
|Engine mass||32 tonne|
|Operating Pressure||500 atm|
|Propellant mass flow||22.3 kg/s|
|Propulsion System||Nuclear Lightbulb|
|Exhaust Velocity||30,000 m/s|
|Specific Impulse||3,058 s|
|Thrust Power||560.7 GW|
|Mass Flow||1,246 kg/s|
|Total Engine Mass||378,000 kg|
|Wet Mass||2,700,000 kg|
|Dry Mass||1,600,000 kg|
|Mass Ratio||1.69 m/s|
|Specific Power||0.67 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.
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.
VCR light bulb fission Exhaust Velocity 19,620 m/s Specific Impulse 2,000 s Thrust 56,400 N Thrust Power 0.6 GW Mass Flow 3 kg/s Total Engine Mass 72,566 kg T/W 0.08 Fuel Fission:
Reactor Gas Core
Remass Seeded Hydrogen Remass Accel Thermal Accel:
Thrust Director Nozzle Specific Power 131 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.
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.
- Nuclear light bulb
- Investigation of gaseous nuclear rocket technology — Summary technical report
- Analytical design and performance studies of nuclear furnace tests of small nuclear light bulb models
- Investigation of gaseous nuclear rocket technology
- Analytical design and performance studies of the nuclear light bulb engine
- Conceptual design studies and experiments related to cavity exhaust systems for nuclear light bulb configurations
- Nuclear light bulb propellant heating simulation using a tungsten/argon aerosol and radiation from a dc arc surrounded by a segmented mirror cavity
- Development of RF plasma simulations of in-reactor tests of small models of the nuclear light bulb fuel region
- Development and tests of small fused silica models of transparent walls for the nuclear light bulb engine
- Experimental investigations to simulate the thermal environment and fuel region in nuclear light bulb reactors using an r-f radiant energy source
- Analytical studies of nuclear light bulb engine radiant heat transfer and performance characteristics
- Experimental investigations to simulate the thermal environment, transparent walls, and propellant heating in a nuclear light bulb engine
- Studies of nuclear light bulb start-up conditions and engine dynamics
- Analytical studies of start-up and dynamic response characteristics of the nuclear light bulb engine
- Nuclear criticality studies of specific nuclear light bulb and open-cycle gaseous nuclear rocket engines
- Nuclear studies of the nuclear light bulb rocket engine
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.
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.
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.
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  cool the exhaust from 6550 K to 1500 K,  prevent the exhaust from severely damaging the exhaust duct,  condense the gaseous uranium into liquid uranium droplets, and  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.
- Fill hydrogen ducts and neon system from storage to a pressure equal to approximately 20 atmospheres
- turn on neon recirculation pump
- inject fuel until critical mass is reached
- increase power level and adjust flow rates and cavity pressure to maintain criticality and limit component temperatures to tolerable level
- inject propellant seeds when 10 percent of full power is reached
- 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.
- Close Fuel Injection Control Valve (turn off the uranium)
- Begin Linear Decrease in Propellant Flow Rate (propellant flow past light bulbs to exhaust nozzles)
- Begin Linear Increase in Radiator Flow Rate (flow from coolant heat exchanger to radiator)
- 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.
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.
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.
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.
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.
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.
|Propulsion System||Gas Core NTR|
|Exhaust Velocity||35,000 m/s|
|Specific Impulse||3,568 s|
|Thrust Power||61.2 GW|
|Mass Flow||100 kg/s|
|Total Engine Mass||200,000 kg|
|Remass Accel||Thermal Accel:|
|Specific Power||3 kg/MW|
|Engine mass||30-200 tonne|
|T/W||11.9 to 1.8|
|Open Cycle 2|
|Propulsion System||Gas Core NTR|
|Exhaust Velocity||50,000 m/s|
|Specific Impulse||5,097 s|
|Thrust Power||0.1 TW|
|Mass Flow||100 kg/s|
|Specific Power||2 kg/MW|
|Engine mass||30-200 tonne|
|T/W||17.0 to 2.5|
|Open Cycle 3|
|Exhaust velocity||25,000 to 69,000 m/s|
|Thrust||19,600 to 108,000 n|
|Engine mass||40 to 110 tonne|
|T/W||0.05 to 0.10|
|Operating Pressure||400 to 2000 atm|
|Open Cycle MAX|
|Exhaust Velocity||98,000 m/s|
|Specific Impulse||9,990 s|
|Thrust Power||0.15 TW|
|Mass Flow||31 kg/s|
|Total Engine Mass||15,000 kg|
|Propulsion System||Gas Core NTR|
|Exhaust Velocity||35,316 m/s|
|Specific Impulse||3,600 s|
|Thrust Power||61.8 GW|
|Mass Flow||99 kg/s|
|Wet Mass||433,000 kg|
|Dry Mass||268,000 kg|
|Mass Ratio||1.62 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.
GAS CORE FISSION THERMAL ROCKETS
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 Thrust
22 750 43.5 52.3 10 6.3 36 10.3 4.3×10-2 44 1500 87 61.6 13 12.6 36 17.5 7.3×10-2 110 3750 218 86 18 32 36 31.3 0.13 220 7500 435 123 24 63 36 43.8 0.18 440 15,000 870 193 31 126 36 55.9 0.23
- 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%.
- 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.
- 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
- 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.
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?
|Exhaust Velocity||15,696 m/s|
|Reactor Mass||2,200 kg|
|Pressure Shell Mass||3,100 kg|
|Radiator Mass||4,930 kg|
|Total Mass||10,230 kg|
|Outer Diameter||1.22 m|
|Cavity Diameter||0.61 m|
|233U Plasma Diameter||0.43 m|
|233U Plasma Mass||1.42 kg|
|233U Plasma Power|
|233U Driver Power|
|Total Engine Power|
|Radiator Alpha||310 kg/MW|
|Exhaust Temp||4,000 to|
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 Mass||<36,000 kg|
|Exhaust Vel||15,700 m/s|
|Exhaust Vel||19,100 m/s|
|Inert Mass||36,000 kg|
|Dry Mass||98,800 kg|
|Propellant Mass||33,620 kg|
|Wet Mass||132,420 kg|
|Exhaust Vel||19,100 m/s|
|Initial Accel||2.27 m/s|
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.
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 Velocity||17,658 m/s|
|Specific Impulse||1,800 s|
|Thrust Power||0.2 TW|
|Mass Flow||1,008 kg/s|
|Total Engine Mass||127,000 kg|
|Remass Accel||Thermal Accel:|
|Specific Power||1 kg/MW|
|Exhaust velocity||22,000 m/s|
|Specific Impulse||2,200 s|
|Engine mass||66,000 kg|
|Fuel Temp||20,000° R|
|Propellant Temp||10,000° R|
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.
Figure 2 shows the mathematical model.
Mass diffusion equation (Schmidt number assumed to be Sc = 0.7):
Intitial and boundary conditions are:
For the region near the inlet:
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.
FUEL VOLUME FRACTION
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 Exhaust Velocity 19,620 m/s Specific Impulse 2,000 s Thrust 50,400 N Thrust Power 0.5 GW Mass Flow 3 kg/s Total Engine Mass 114,116 kg T/W 0.04 Frozen Flow eff. 75% Thermal eff. 70% Total eff. 53% Fuel Fission:
Reactor Gas Core
Remass Seeded Hydrogen Remass Accel Thermal Accel:
Thrust Director Nozzle Specific Power 231 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.
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.
mass flow ratio
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 bimodal GCNTR|
|Specific Impulse||2,000 sec|
|Chamber Pressure||500 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
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 http://www.npoenergomash.ru/dejatelnost/engines/, 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 (https://ru.wikipedia.org/wiki/%D0%A0%D0%94-0410) that contained a timeline apparently copied from https://mipt.ru/education/chairs/physmech/science/research/nuclear_reactor.php, which is a webpage of the Department of physical mechanics of the Moscow physico-technical university. Translation follows:
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.
Backtracking a bit, I’ve found an earlier memo containing some of the points missing from the above report
Here’s the economic aspect:
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.
And finally, here’s an essay from a defunct… site, credited to an Aleksandr Valeryevich Khoroshikh at horoshih-aleksander at yandex.ru); it has images, citations, technical details, and it offers a saddening finale to the whole story.
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 http://militera.lib.ru/explo/chertok_be/index.html. №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 http://novosti-kosmonavtiki.ru/mag/2001/1411/ and http://novosti-kosmonavtiki.ru/mag/2005/1043/.
№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 http://engine.aviaport.ru/issues/05/page41.html and http://engine.aviaport.ru/issues/06/page12.html. Aside from several more images, I’ve gleaned two very interesting sections not covered above:
Estimated parameters of a gas-core fuel element Pressure in reaction chamber, kgf/cm2 200 Uranium expenditure, g/sec 200 Hydrogen expenditure in the reaction chamber, g/sec 10 Velocity of fuel when entering the reaction chamber, m/s 1.7 Power, kWt 1,000 Share of vaporized uranium in the egress flow, % 80 Temperature of uranium plasma, K 8—10×103(unclear – ed.) Thermal neutron flow, neutrons/cm2/sec 1015(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,
2% UTB solution
|Exhaust Velocity||66,000 m/s|
|Specific Impulse||6,728 s|
|Thrust Power||425.7 GW|
|Mass Flow||195 kg/s|
|Plenum Radius||3.075 cm|
|Plenum Length||65 cm|
|Total Engine Mass||33,000 kg|
|Remass Accel||Thermal Accel:|
|Thrust Director||Pusher Plate|
|Specific Power||77.5 kg/GW|
(missing items same as above)
|Exhaust Velocity||4,725,000 m/s|
|Specific Impulse||482,140 s|
|Thrust Power||30,600 GW|
|Mass Flow||3 kg/s|
|Specific Power||1.1 kg/GW|
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:
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 vΣf-Σa)/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...
Zubrin NSWR Exhaust Velocity 78,480 m/s Specific Impulse 8,000 s Thrust 8,696,900 N Thrust Power 0.3 TW Mass Flow 111 kg/s Total Engine Mass 495,467 kg T/W 1.79 Frozen Flow eff. 80% Total eff. 80% Fuel Fission:
Reactor Gas Core
Remass Water Remass Accel Thermal Accel:
Thrust Director Pusher Plate Specific Power 1.45 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).
My friend Robert Zubrin published a paper about Nuclear Salt Water Rockets back in 1991.
Here, you took water along with nuclear salts, and stored them in neutron absorbing tanks, exhausting the water through a nozzle that allowed the salts to achieve criticality. You could achieve very high exhaust velocities that way!
Another approach is possible. Namely, putting Lithium-6 Deuteride in water, along with neutron multiplying materials, and exposing the fusion fuels to a high neutron flux in the rocket engine. Similar exhaust velocities can be achieved. Without the radioactive byproducts of the Nuclear Salt Water Rocket approach.
There are several advantages relative to conventional NTR designs. As the peak neutron flux and fission reaction rates occur outside the vehicle, these are far greater than what is possible when built into the vessel. A contained reactor can only allow a small percentage of its fuel to undergo fission at any given time, otherwise it would overheat and meltdown or explode in a runaway fission chain reaction. The fission reaction in an NSWR is dynamic and because the reaction products are exhausted into space it doesn't have a limit on the proportion of fission fuel that reacts.
NSWRs are a hybrid between fission reactors and fission bombs.
Due to their ability to harness the power of what is essentially a continuous nuclear fission explosion, NSWRs would have both very high thrust and very high exhaust velocity. The rocket would be able to accelerate quickly as well as be extremely efficient in terms of propellant usage. Zubrin proposed one design that generates 13 meganewtons of thrust at 66 km/s exhaust velocity. Another design achieves 4,700 km/s and uses 2,700 tonnes of highly enriched uranium salts in water to propel a 300 tonne spacecraft up to 3.6% of the speed of light.
This basically solves the problem of spaceflight when done with Lithium-6 Deuterium Jetter Cycle process.
That cycle may be sustained by a high flux of neutrons in the engine core. The High Flux Isotope Reactor is a model for the type of engine I'm talking about.
Here neutrons are focused into a central tube through which water passes. The water moderates the neutrons passing through it. A beryllium reflector keeps the neutrons in the reactor. Lithium-6 Deuteride suspended in the water absorbs the neutrons and supports Jetter Cycle fusion. Helium gas, neutrons and steam are the only exhaust products.
NSWRs share many of the features of Orion propulsion systems, except that NSWRs generate continuous thrust and work on much smaller scales than the smallest feasible Orion designs.
A single stage system I described previously, with one crewman and five passengers seated in a capsule beneath a 30 cubic meter propellant tank, would mass 1,585 kg and carry 30,000 kg of water salted with 6LiD (mass ratio of 19.9), passing through a high neutron flux region to produce controlled thrust. With a 4,700 km/sec exhaust speed, the vehicle is capable of achieving (a delta-V of) 14,062.86 km/sec! Enough for a one gee boost of 16.6 days!!
With 4 days of boost combined with 4 days of slowing down the ship can cruise at one gee a distance of 1,171.28 million km (torchship brachistochrone trajectory). This is sufficient to fly to any celestial body out to Jupiter and back. Reducing acceleration after planetary escape to 0.416 gees increases boost time to 10 days per leg, 40 days per round trip, and increases range to 6,075.15 million km. This is sufficient to take us to all celestial bodies in the solar system, including Pluto, Haumea, Makemake, and Eris.
One gee Mercury 57.9 91.7 207.5 2.24 3.37 167.9 3.03 Venus 108.2 41.4 257.8 1.50 3.75 151.1 2.87 Earth 149.6 Mars 227.9 78.3 377.5 2.07 4.54 170.6 3.05 Jupiter 778.4 628.8 928.0 5.86 7.12 939.8 7.17 0.42 gee Saturn 1,428.7 1,279.1 1,578.3 12.98 14.42 1,437 13.76 Uranus 2,871.0 2,721.4 3,020.6 18.93 19.94 2,952 19.72 Neptune 4,495.3 4,345.7 4,644.9 23.92 24.73 4,360 23.96
Building a ship such as this, and flying it back to the moon, would be an appropriate opening to the second round of the space age. In less than a month such a ship could do a 'grand tour' of all the planets out to Jupiter.
On 24 July 2014 The Grand Tour would be:
Grand Tour Dist Days Earth⇒Mars 170.6 3.05 Mars⇒Venus 330.7 4.25 Venus⇒Mercury 64.8 1.88 Mercury⇒Jupiter 780.3 10.14 Jupiter⇒Earth 939.8 11.12 TOTAL 2286.20 30.44
The hop from Mercury to Jupiter and back to Earth would occur at 41% earth normal gravity to conserve fuel.
A hop from Earth to moon would take 3.75 hours! Then a hop to Mars, taking 3.05 days. Then off to Venus 4.25 days. Thence to Mercury, 1.88 days. Then the long haul to Jupiter 10.14 days at reduced gravity. Then back to Earth again at reduced gravity 11.12 days.
The five passengers brought along would pay $100 million each — which would pay for the entire development programme. The promotion of the power source used in the rocket, which would be developed along with the engine, would also pay dividends for the lucky passengers.
The reaction conditions for a power plant is quite different than that for a rocket exhaust.
A 1000 MW steam turbine for the US power plant Ravenswood Unit 3 was cross-compound with 16.6 Mpa, 538C steam conditions with 44.7% thermal efficiency. These steam conditions give the density of Li-6 Deuteride needed for a given neutron flux to maintain 1000 MW electrical output. A closed cycle steam system can be quite compact for this unit, with no exhaust to the atmosphere except helium, but even that can be mined from the closed cycle system. Lithium-6 Deuteride is added to the water tank to maintain power levels.
1000 MW operating continuously and selling power for $0.10 per kWh produces $876.6 million per year revenue. Discounted at 6.25% per year over 30 years this revenue stream is worth $11.75 billion. Profits of $10 billion are possible for each unit sold. Half these profits from the first unit divided among the passengers, translates to $1 billion returns for the first power plant installed — with 50% revenue flowing to the money investors — and 50% flowing to the company — so, the trip plus $1 billion is what each passenger gets for their $100 million. They can even sell forward their seat as the vehicle nears flight readiness, and increase their revenue.
This small ship described here massing 31.6 metric tons at lift off, with a take off acceleration of 2 gees, requires a mass flow rate of 142 grams per second with an exhaust velocity of 4,300 km/sec. This ship during lift-off produces an exhaust jet power of 1.31 trillion watts! The ability to reliably sustain this level of power safely and reliably in a compact space, is an obvious promotion of the company's capacity to build reliable compact and safe nuclear power units. Five 1000 MW turbines installed each year generate $50 billion in free cash flow, when monetized at discount rates.
What do you with the money? Build bigger ships!
The ships scale to any size. You open the solar system with them, and depopulate the Earth.
|Specific Impulse||1,000,000 sec|
|Exhaust velocity||9,810,000 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.
|Thrust Power||0.2 GW|
|Mass Flow||1.00e-06 kg/s|
|Specific Power||55 kg/MW|
|Thrust Power||2.6 GW|
|Mass Flow||2.30e-05 kg/s|
|Specific Power||3 kg/MW|
|Exhaust Velocity||15,000,000 m/s|
|Specific Impulse||1,529,052 s|
|Total Engine Mass||9,000 kg|
|Thrust Director||Magnetic 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.
|Exhaust Velocity||5,170,000 m/s|
|Specific Impulse||527,013 s|
|Thrust Power||0.1 GW|
|Mass Flow||8.00e-06 kg/s|
|Total Engine Mass||113,400 kg|
|Thrust Director||Magnetic Nozzle|
|Specific Power||1,020 kg/MW|
|Propulsion System||Werka FFRE|
|Wet Mass||303,000 kg|
|Dry Mass||295,000 kg|
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.
|Reactor Power||2.5 GW|
|Thrust Power||730 MW|
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.
|FFRE||527,000 sec||43 Newtons|
|AFFRE||32,000 sec||4,651 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.
|Antimatter Sail to 240 AU|
|Trip Distance||240 AU|
|Trip Time||10 years|
|Mission ΔV||116,800 m/s|
|Exhaust Velocity||6.7×104 m/s|
|Electrical output||400 We|
|Specific Mass||16 kg/kW|
|Carbon Sail||0.7 kg|
|Power Plant||6.4 kg|
|Antmatter Storage||9 kg|
|Inert Mass||16 kg|
|Dry Mass||26 kg|
|Uranium fuel||109 kg|
|Wet Mass||135 kg|
|Initial Accel||0.006 m/s2|
This is from NIAC Phase I Progress Report Antimatter Driven Sail for Deep Space Missions (2004)
They initially wanted to design a spacecraft that could transport a few kilograms of instrument package to another star system. Since that was a daunting goal, they decided to do baby-steps first and design for the less demanding mission of sending a payload 250 AUs to the Kuiper Belt in ten years (preferably braking to a halt but fly-by is acceptable). Even this is way beyond the current state of the art.
Their final design can send a 10 kilogram instrument package 250 AUs in 10 years using 30 milligrams of antimatter. That is a lot of antimatter, but the US could produce it in as little as 40 years with current particle accelerators. The average velocity was figured to be 116.8 km/s.
The design could send 10 kg of instruments to Alpha Centauri in 40 years using 17 grams of antimatter. A previous JPL study had concluded that kilograms of antimatter would be needed, the sail design uses only a fraction of that. Of course 17 gm of antimatter would take about 23,000 years to produce with current particle accelerators, more efficient antimatter production methods would be needed.
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.
The critical factors are:
- The momentum transfered to the sail from the antimatter-induced fission has to be maximized (if the fission does not give the sail enough thrust you'll need outrageous amounts of antimatter)
- Technology needs to be developed to store anti-hydrogen micro-pellets within solid-state integrated circuits (antimatter+matter goes boom! Antimatter containment system cannot weigh tons, it has to be very low mass)
- Technology needs to be developed to create a high specific-mass electrical-power supply based on antimatter fission conversion (power supply cannot weigh tons either. Neither can the fuel supply)
The critical factor is "Nat", or number of uranium atoms ejected per fission. Nat has a remarkably strong impact on the number of antiprotons required for the mission.
The carbon-carbon sail contains uranium. A uranium atom undergoes fission when impacted by an antiproton. If only two fission fragments are released, then the momentum is determined by the velocity and mass of one of the fragments.
The fission fragments vary in mass, but they can be approximated by assuming the average fission product is palladium-111, mass is 1.8×10-25 kg/atom. The fission releases 190 MeV, which would give the palladium-111 atom an exhaust velocity of 1.39×107 m/s. This corresponds to a specific impulse of around 1.4 million seconds (i.e., dividing by 9.81).
Now, if neutral atoms of uranium are blown off with each fission event in addition to the palladium-111 fragments, the energy is distributed to all the ejected particles. Increasing the number of neutral uranium atoms (increasing Nat) will increase the thrust on the sail, at the cost of reducing the specific impulse. Yes, it's our old friend "shifting gears" again.
This is important. If you have a fixed-thrust rocket engine, and a given deltaV for the mission, you can calculate the optimal exhaust velocity and specific impulse. For our 116.8 km/s Kuiper mission this come out to about an exhaust velocity of 6.7×104 m/s (Isp of 6,800 seconds).
So by controlling Nat, one can control the exhaust velocity.
Nat is controlled by altering the velocity of the antiprotons, in other words the antiproton beam energy. The faster they go, the deeper they penetrate the uranium layer, and the more uranium atoms are ejected by each fission event. This is done by electrostatically biasing the sail relative to the antihydrogen container. This can easily raise the antiproton beam energy to 100 keV, which will allow them to penetrate the uranium to a depth of 355 nanometers below the surface.
As Nat increases, so does the consumption of uranium fuel. For the Kuiper belt mission, the uranium layer will have to have a thickness of 293 microns.
The researchers did a parametric study using Nat (number of atoms ejected per fission) as a free parameters.
Figure 1 shows the specific impulse as a function of Nat. We have assumed that the energy released in fission is equally distributed among the atoms ejected.
Figure 2 shows the dependence of the mass of antimatter required to perform the mission. The figure clearly depicts a minimum for Nat equal to around 15,000. This corresponds to a specific impulse of almost 7500 sec.
Figure 3 shows the masses of the uranium fuel and the spacecraft as a function of Nat. For Nat equal to 1, the specific impulse is over a million seconds and the fuel mass is small compared to the ship mass. As the specific impulse gets below 10,000 s, the mass of the fuel begins to dominate the ship mass.
For the Kuiper mission, the optimum value for Nat, i.e. the value where the number of required antihydrogen atoms was a minimum, occurred at Nat = 15,000 atoms per fission. So this value was used for the system studies.
The sail is 5 meters in diameter. It composed of a carbon-carbon layer 15 microns thick (34 g/m2), coated with a uranium fuel layer 293 microns thick. The total mass of uranium fuel is 109 kilograms.
Since the sail is so thin it does not need any heat radiators. It can dissipate the waste heat passively by black-body emmission. Assuming an emissivity of 0.3 for the carbon and uranium, the steady state temperature would be 570° C, well below uranium's melting point.
Antihydrogen storage unit
The antimatter storage is held 12 meters away from the sail by four tethers.
The storage is an array of small chips that resemble integrated circuit chips but are not. They are a series of tunnels etched in a silicon subtrate. Each tunnel is a sequence of electrodes. Each pair of electrodes holds a tiny pellet of antimatter fuel, in the form of solid antihydrogen. Each pellet has about 1015 antihydrogen atoms, charged to about 10-11 coulombs.
Each tunnel has 67 electrode pairs holding 67 antihydrogen pellets. Each 4 cm long chip has 100 tunnels, so each chip contains 1.6×1019 antihydrogen atoms. There are roughly 2,000 chips in the storage unit. Total number of antihydrogen atoms is 1.8×1022 or 30.45 milligrams. The entire storage unit has a mass of about 9 kilograms.
As pellets are ejected toward the sail, the remaining pellets in a tunnel are shuffled towards the sail-end like a bucket brigade.
My back of the envelope calculation says that if the entire storage array failed the resulting blast will be roughly the same as a 93 kiloton nuclear bomb, a little less than a W76 thermonuclear warhead.
The spacecraft requires electricity. Since a nuclear reactor or RTG is too heavy, and solar power is pretty worthess when you get further from the Sun than the asteroid belt, something new is needed. Since the ship is already carrying the most powerful fuel in the universe, the study authors tried to design an Antimatter Fission Conversion (AFC) power plant.
Some of the antimatter storage tunnels point rearward at the AFC uranium cone. The antimatter causes uranium atoms to fission. The scintillator converts the moving fission fragments into light (finding a scintillator material with the proper properties will be difficult). The photovoltaic cells convert the light into electricity. The AFC has a liquid lithium heat exchanger that removes the waste heat to a beryllium heat radiator, otherwise the blasted AFC would melt.
The spacecraft power requirements were estimated as 400 watts (same as the Voyager spacecraft). Voyager was powered by three MHW-RTG with a combined mass of 113.1 kg. The AFC will have a mass of only 6.4 kg, specific mass of 16 kg/kW (6.4 / 0.4).
The over-all efficiency of the AFC was estimated to be about 4.4%, so 2×1014 antiprotons per second are needed for 400 W of electrical power. The power is generated on demand when communications back to Terra are indicated. The beryllium heat radiator is in two dametrically opposed sheets, edge-on to the sail. They have a radiating surface area of 3.5 m2 and a temperature of 620° C.
The researchers studied the effect of increasing the spacecraft's dry mass (power plant, structure, payload, etc). The result is in Figure 7. Mpbar (mg) is the mass of antimatter fuel in milligrams. Mu (kg) is the mass of uranium fuel in kilograms. Mship (kg) is the wet mass of the spacecraft in kilograms. Delta dry mass (kg) scale is the increase in the ship's dry mass (baseline ship has a dry mass of 26 kg). Mass scale is in kilograms or milligrams depending upon the weight unit in parenthesis.
So if the dry mass is doubled (delta dry mass = 26 kg, total dry mass = 52 kg) then mass of ship only rises to Mship = 280 kg and the amount of antimatter fuel rises to Mpbar = 60 mg.
The baseline ship is where delta dry mass = 0 kg. Mpbar = 30.45 milligrams, Mu = 109 kilograms, Mship = 135 kg.
The minimum acceleration level was decided to be 0.006 m/s2. If my slide rule is not lying to me, that implies a thrust of 0.81 Newtons.
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.