## Introduction

There is a nice basic overview of propulsion systems here.

You can spend lots of time researching spacecraft propulsion systems. But you are in luck, I've got some data for you. This is from various places I found around the internet, and no, I didn't keep track of where I got them. Use at your own risk.

If you don't like the values in the table, do some research to see if you can discover values you like better. Also note that the designs in the list are probably optimized for high exhaust velocities at the expense of thrust. There is a chance that some can be altered to give enough thrust for lift-off at the expense of exhaust velocity. Or you can just give up and go beg Mr. Tyco Bass for some atomic tri-tetramethylbenzacarbonethylene. Four drops should do the trick.

Some engines require electricity in order to operate. These have their megawatt requirements listed under "Power Requirements". With these engines, the Engine Mass value includes the mass of the power plant (unless the value includes "+pp", which means the mass value does NOT include the mass of the power plant). The power plant mass can be omitted if the spacecraft relies on beamed power from a remote power station. Alas, I could find no figures on the mass of the power plant. If the plant is nuclear, it probably has a mass of around 0.5 to 10 tons per megawatt. If it is beamed power the mass is of course zero. Efficiency is the percentage of the power requirements megawatts that are actually turned into thrust. The rest becomes waste heat and has to be removed with heat radiators.

T/W >1.0 = Thrust to Weight ratio greater than one? This boils down to: can this engine be used to take off from Terra's surface? If the answer is "no" use it only for orbit to orbit maneuvers. It is calculated by figuring if the given thrust can accelerate the engine mass greater than one gee of acceleration. As a general rule, a practical spacecraft capable of lifting off from the Earth's surface will require a T/W of about 50 to 75.

Most propulsion systems fall into two categories: SUV and economy. SUV propulsion is like an SUV automobile: big and muscular, but the blasted thing gets a pathetic three miles to the gallon. Economy propulsion has fantastic fuel economy, but has trouble climbing low hills. In the world of rockets, good fuel economy means a high "specific impulse" (Isp) and high exhaust velocity. And muscle means a high thrust.

The technical terms for SUV high-thrust + low-specific-impulse are Specific-Impulse Limited and High-Thrust Systems. Typical examples are chemical and solid-core nuclear thermal. These usually create the exhaust velocity by thermal means (heat), so they are limited by how hot you can get the exhaust (limited by chemical energy or limited by the melting point of the rocket engine).

The technical terms for Economy low-thrust + high-specific-impulse are Specific-Mass Limited, Low-Thrust Systems, and Power-Limited Systems. Typical example is an ion drive. "Specific Mass" or "Alpha" (α) is the mass of the propulsion system divided by the thrust power. These are usually electrically powered rockets, which is why they are power-limited.

The only vaguely possible propulsion system that has both high exhaust velocity and high thrust is the Nuclear Salt Water Rocket, and not a few scientist have questions about its feasibility. Well, actually there is also Project Orion, but that has other problems (see below). In science fiction, one often encounters the legendary "fusion drive" or "torchship", which is a high exhaust velocity + high thrust propulsion system that modern science isn't sure is even possible.

## The Drive Table

Note that this table only contains engines for which I have data for the engine's thrust. There are a few for which I only have the specific impulse (e.g., Positron Ablative, LH2/Fluorine, Photon, etc.). These do not appear on the table but they have entries below.

All drives listed in the table whose names end in "MAX" require some sort of technological breakthrough to to prevent the engine from vaporizing and/or absurdly large reaction chamber sizes.

If these figures result in disappointing rocket performance, in the name of science fiction you can tweak some of them and claim it was due to a technological advance. You are allowed to tweak anything who's name does not end in "MAX". You can alter the Thrust, Engine Mass, and/or the Eff, but no other values. If there is a corresponding "MAX" entry for the engine you are tweaking, you cannot alter any of the values above the "MAX" entry (i.e., you are not allowed to tweak NTR-SOLID-DUMBO's thrust above 7,000,000, which is the value in the NTR-SOLID MAX entry).

The engines are sorted by thrust power, since that depends on both exhaust velocity and thrust. So engines that high in both of those parameters will be towards the end of the list. This is useful for designers trying to make spacecraft that can both blast-off from a planet's surface and do efficient orbital transfers.

As Philip Eklund noted in his game High Frontier, the engines fall into three rough categories: megawatt thrusters (thrust power), gigawatt thrusters, and terawatt thrusters. Though if you want to be pedantic the radioisotope, ArcJet, and HOPE MPD engines are kilowatt thrusters. The resistojet is a hectowatt thruster and the poor little DAWN NSTAR is a pathetic watt thruster.

If one was trying to design a more reasonable strictly orbit-to-orbit spacecraft one would want the engine list sorted by exhaust velocity. And surface-to-orbit designers would want the list sorted by thrust. Sorry, you'll have to do that yourself.

PropulsionCodeThrust
Power
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
DAWN mission NSTARESTAT1.37 W30,4113,1009.00e-05263.60e-07
ResistojetETHERM725 W2,9002961
ArcJetETHERM20 kW20,0002,0392
HOPE Cargo MPDEMAG432 kW78,5008,00211
HOPE Tanker MPDEMAG432 kW78,5008,00211
START OF MEGAWATT THRUSTERS
PropulsionCodeThrust
Power
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
HOPE Crew MPDEMAG1.1 MW78,5008,00228
Magneto Inertial Fusion (low)PULSE2.6 MW50,4205,140103
VASIMR (low gear)EMAG5.9 MW29,0002,95640010,0000.004
VASIMR (high gear)EMAG5.9 MW294,00029,9694010,0004.08e-04
VASIMR (med gear)EMAG5.8 MW147,00014,9858010,0008.15e-04
Space Shuttle RCSCHEM LIQ6.0 MW3,1003163,8704106.620
Mirror SteamerBEAM12.8 MW9,8101,0002,60020,9770.013
Monatomic-H MITEENTR SOLID15.0 MW12,7501,3002,3502001.198
HybridMITEEETHERM15.0 MW17,6601,8001,70010,0000.017
PropulsionCodeThrust
Power
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
AIMPULSE16.5 MW598,00060,95855
Solar MothBEAM18.0 MW9,0009174,0001004.077
Antimatter-Driven SailNTR FRAG19.5 MW73,6007,5005301090.496
Umbrella ShipESTAT19.7 MW80,4428,200490
ArcJetETHERM31.4 MW19,6202,0003,20022,3690.015
Hall EffectESTAT32.4 MW19,6202,0003,30085,4690.004
Magneto Inertial FusionPULSE36.0 MW49,0004,9951,4709,1000.016
Pulsed Plasmoid ThrusterPULSE43.2 MW78,4808,0001,10083,6110.001
Ponderomotive VASIMREMAG44.2 MW39,2404,0002,25043,7960.005
Wakefield E-BeamETHERM45.1 MW19,6202,0004,60041,8370.011
Ablative LaserBEAM47.1 MW39,2404,0002,40022,2220.011
MPD T-WaveEMAG47.1 MW78,4808,0001,20082,6750.001
Tungsten ResistojetETHERM48.6 MW9,8101,0009,90042,6010.024
NASA space tugCHEM LIQ49.3 MW4,40044922,400199,6000.011
Mass DriverOTHER51.0 MW9,8101,00010,400163,0000.007
Ion DriveESTAT56.7 MW78,4808,0001,444120,1490.001
MET Steamer AmplitronsETHERM58.9 MW9,8101,00012,000123,3020.010
NERVA (CO or N2)NTR SOLID64.9 MW2,64927049,00010,0000.499
LPNTR High GearNTR SOLID65 MW13,2001,3509,8008351.2
Basic MITEENTR SOLID68.7 MW9,8101,00014,0002007.136
NERVA (CO2)NTR SOLID81.0 MW3,30633749,00010,0000.499
NERVA (H2O)NTR SOLID99.0 MW4,04241249,00010,0000.499
HOPE FFRENTR FRAG111.2 MW5,170,000527,01343
Werka FFRENTR FRAG111.2 MW5,170,000527,01343113,4003.90e-05
NERVA (NH3)NTR SOLID125.0 MW5,10152049,00010,0000.499
D-D Fusion InertialPULSE125.6 MW78,4808,0003,200243,3330.001
NERVA (CH4)NTR SOLID154.8 MW6,31864449,00010,0000.499
Twisted Ribbon RD-0140NTR SOLID155.7 MW8,83090035,2802,0001.798
Dusty Plasma (550AU)NTR FRAG165.0 MW15,000,0001,529,052229,0002.49e-04
PropulsionCodeThrust
Power
(MW)
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
Colloid ThrusterESTAT17243,0004,3838,00020,0000.041
LARSNTR LIQUID19619,6202,00020,0001,0002.039
NERVA (H2)NTR SOLID1988,09382549,00010,0000.499
Laser ThermalBEAM26040,0004,07713,00020,0000.066
LPNTR Low GearNTR SOLID29211,9001,21049,0008356.0
Mass DriverOTHER30030,0003,05820,000150,0000.014
Twisted Ribbon NPPSNTR SOLID3079,02492068,0001,8003.851
LighterCHEM LIQ3094,410450140,000
LANTR (high gear)NTR SOLID3099,22194067,000
SNRE-classNTR SOLID3228,83090073,0002,4003.010
Magneto Inertial Fusion (hi)PULSE34850,4205,14013,800
H-B Cat InertialPULSE369156,96016,0004,70065,0890.007
Aluminum/LOX rocketCHEM LIQ3882,649270292,60056,0000.533
NERVA (H)NTR SOLID39216,0001,63149,00010,0000.499
MICF Fusion (cDD 100/sec)FUSION42336,3003,70023,300
Kuck MosquitoCHEM LIQ4844,400449220,000
Vortex Confined (H2)NTR GAS OP49419,6202,00050,400114,1160.045
Pewee-classNTR SOLID511.59,200938111,2003,2403.499
LH2/LOX rocketCHEM LIQ5404,905500220,00026,6670.841
VCR Light Bulb (H2)NTR GAS CL55319,6202,00056,40072,5660.079
LANTR (low gear)NTR SOLID5846,347647184,000
Dual-mode Fission (H2)NTR SOLID6129,8101,000124,70033,0000.385
H-Li6 FusorFUSION658.319,6202,00067,10054,0000.127
Cermet NERVA (H2)NTR SOLID6599,8101,000134,40032,5460.421
Z-Pinch MicrofissionPULSE667156,96016,0008,500193,3330.004
Afterburner FFRENTR FRAG730313,90032,0004,651268,9610.002
Dual-mode PB (H2)NTR SOLID8479,8101,000172,70058,0000.304
Bimodal NTR Solid (NASA)NTR SOLID8988,980915200,0006,6723.056
MICF Fusion (DT 100/sec)FUSION91144,2004,50041,200
START OF GIGAWATT THRUSTERS
PropulsionCodeThrust
Power
(MW)
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
IonESTAT1,050210,00021,40710,000400,0000.003
D-T FusionFUSION1,18822,0002,243108,00010,0001.101
NERVA Deriv (H2)NTR SOLID1,3508,085824334,06110,1003.372
Antimatter BottlePULSE1,36278,4808,00034,700180,0000.020
Metastable He*CHEM1,37643,0004,38364,00010,0000.652
Resuable Nuclear ShuttleNTR SOLID1,3768,000815344,000
Metastable HeliumCHEM1,56729,4303,000106,500
n-6Li MicrofissionPULSE1,570156,96016,00020,000106,6670.019
Pebble Bed (H2)NTR SOLID1,5909,530971333,6171,70020.005
Vapor Core (H2)NTR VAPOR1,6179,800999330,0006,8304.925
Twisted ribbon NTRNTR SOLID1,6509,420960330,0005,2606.4
3He-D Mirror CellFUSION1,664313,92032,00010,600106,6670.010
HOPE MTFFUSION2,005691,46070,4905,798116,0210.005
Cermet (H2)NTR SOLID2,0309,120930445,2679,0005.043
D-T Fusion TokamakFUSION2,23166,8006,80966,800197,0000.035
Widmer Mars MissionNTR SOLID2,3208,000815580,000
p-Nerva (NRX)AM SOLID2,37010,7901,100440,00010,9404.100
Antimatter Solid maxAM SOLID2,37410,7911,100440,000
Dusty Plasma (0.5LY)NTR FRAG2,58015,000,0001,529,0523449,0000.004
Proton RD-253 x1CHEM LIQ2,8363,1003161,830,0001,260148.051
PuFF Pulsed Fission Fusion2,880196,00019,98029,40055,5600.054
Discovery II (MC Toroidal)FUSION3,123347,00035,37218,000
HOPE Z-PinchPULSE3,617189,78019,34638,12095,1380.041
HELIOS 2nd StageNTR SOLID3,8267,800795981,000
NTR Gas/Closed (H2)NTR GAS CL4,54020,4052,080445,00056,8000.799
Dumbo (CO or N2)NTR SOLID4,6362,6492703,500,0005,00071.356
Atomic V-2NTR SOLID4,7148,9809151,050,0004,20025.484
Fusion Asteroid MinerFUSION4,800100,00010,19048,00010,0000.979
ORION FissionPULSE5,65443,0004,383263,000200,0000.134
Dumbo (CO2)NTR SOLID5,7863,3063373,500,0005,00071.356
THS Fusion Pulse high gearPULSE6,000300,00030,58140,0004,0001.019
THS Fusion Pulse low gearPULSE6,000150,00015,29180,0004,0002.039
PropulsionCodeThrust
Power
(MW)
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
Dumbo (H2O)NTR SOLID7,0744,0424123,500,0005,00071.356
Antiproton SailNTR FRAG8,05930,411,0003,100,000530
Dumbo (NH3)NTR SOLID8,9275,1015203,500,0005,00071.356
ORION FusionPULSE10,65873,0007,441292,000200,0000.149
Dumbo (CH4)NTR SOLID11,0566,3186443,500,0005,00071.356
Saturn-V F-1 x1CHEM LIQ11,5412,9823047,740,5009,15386.206
ACMF (ICAN-II)PULSE11,919132,43513,500180,00027,0000.680
Space Shuttle SSME x3CHEM LIQ12,1154,4444535,452,2009,53158.313
Dumbo (H2)NTR SOLID14,1638,0938253,500,0005,00071.356
Solid rocketCHEM SOLID15,0003,00030610,000,000
Proton RD-253 x6CHEM LIQ16,2283,10031610,470,0007,560141.174
VISTAPULSE20,400170,00017,329240,000
Project OrionPULSE21,73119,6202,0002,215,200203,6801.109
Nuclear DC-X NERVANTR SOLID27,2729,8101,0005,560,000199,6002.840
Dumbo (H)NTR SOLID28,00016,0001,6313,500,0005,00071.356
Mini-Mag Orion (DRM-3)PULSE29,85393,0009,480642,000199,6000.328
Antimatter Plasma (H2O)AM PLASMA29,890980,00099,89861,000500,0000.012
H-B FusionFUSION29,890980,00099,89861,000300,0000.021
Mini-Mag Orion (DRM-1)PULSE29,90693,1649,497642,000119,0460.550
APCP Space Shuttle SRB x2CHEM SOLID31,2002,60026524,000,0001,180,0002.073
ORION USAF 10mPULSE32,90032,9003,3542,000,000107,9001.889
NTR Solid MAXNTR SOLID42,00012,0001,2237,000,00015,00047.571
Gasdynamic MirrorFUSION46,0601,960,000199,79647,000
Nuclear DC-X LANTRNTR SOLID49,2065,90060116,680,000199,6008.519
NTR Liquid maxNTR GAS OP56,00016,0001,6317,000,00070,00010.194
LunaNTR LIQUID56,65010,3001,05011,000,0009,000124.589
PropulsionCodeThrust
Power
(MW)
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
Saturn-V F-1 x5CHEM LIQ58,0543,00030638,702,50045,76586.206
NTR Gas/Open (H2)NTR GAS OP61,25035,0003,5683,500,000200,0001.784
TankerNTR GAS OP61,80335,3163,6003,500,000
AV:T high gearPULSE102,138832,92884,906245,250
NTR Gas/Open 2nd GenNTR GAS OP125,00050,0005,0975,000,000200,0002.548
Mini-Mag OrionPULSE146,795157,00016,0041,870,000199,6000.955
NTR Gas MAXNTR GAS OP147,00098,0009,9903,000,00015,00020.387
NTR Gas/Coaxial (H2)NTR GAS OP157,15617,6581,80017,800,000127,00014.287
Antimatter Plasma (H2)AM PLASMA192,0807,840,000799,18549,000500,0000.010
He3-D FusionFUSION192,0807,840,000799,18549,0001,200,0000.004
MC-Fusion MAXFUSION200,0008,000,000815,49450,0006008.495
P-jet MagnetoInertial (con)FUSION309,3501,189,790121,280520,000
Salt-water ZubrinNTR GAS OP341,26678,4808,0008,696,900495,4671.789
NSWR (20% UTB)NTR GAS OP425,70066,0006,72812,900,00033,00039.848
Liberty ShipNTR GAS CL560,70030,0003,05837,380,000378,00010.080
Cargo Tug SlingshotESTAT764,400280,00028,5425,460,000
Start Of Terawatt Thrusters
PropulsionCodeThrust
Power
(MW)
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
IBS AgamemnonESTAT1,100,000220,00022,42610,000,000
ORION battleshipPULSE1,560,00039,0003,97680,000,0001,700,0004.797
AV:T low gearPULSE2,541,895104,11610,61348,828,125
P-jet MagnetoInertial (med)FUSION2,989,0514,461,270454,7701,340,000
Ghost StarshipFUSION4,237,9205,297,400540,0001,600,000
Epstein DriveFUSION5,500,00011,000,0001,100,0001,000,000
P-jet MagnetoInertial (opt)FUSION7,941,0008,922,480909,5301,780,000
Firefly Starship (Z-Pinch)FUSION12,245,50012,890,0001,313,9701,900,000
NSWR (90% UTB) MAXNTR GAS OP30,550,0004,700,000479,10313,000,000
ORION MAXPULSE39,200,0009,800,000998,9818,000,0008,000101.937
Enzmann StarshipFUSION343,570,50011,700,0001,192,70058,730,000
IC-Fusion MAXPULSE500,000,00010,000,0001,019,368100,000,0001,000,00010.194
Antimatter Beam MAXAM BEAM500,000,000100,000,00010,193,68010,000,00010,000101.937
Frisbee StarshipAM BEAM586,413,00099,900,00010,183,48611,740,000

## Antimatter

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

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

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

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

### Antimatter Energy

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

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

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

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

Charged Pions

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

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

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

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

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

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

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

Gamma Rays

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

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

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

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

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

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

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

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

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

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

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

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

Eγ = (Eπ± * 1.506) - Eπ±

where:

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

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

Eγ = (Eπ± * 1.506) - Eπ±

Eγ = (200 * 1.506) - 200

Eγ = 101.2 megawatts of lethal gamma rays

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

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

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

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

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

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

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

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

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

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

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

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

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

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

### Solid Core

p-Nerva engine (NRX)
Thrust4.4×105 N
Thrust Power2.7 GW
Engine Mass11,000 kg
T/W4.1
T/W >1.0yes
Specific Impulse1,100 sec to
1,300 sec
Exhaust Velocity10,790 m/s to
12,750 m/s
Fuelantiprotons (p)
Fuel Mass Flow13 μg/sec
(1.3×10-8 kg/s)
PropellantLH2
Propellant
Mass Flow
40.7 kg/s
p-LH2 Mix1×10-6 kg p per
7,000 kg LH2
Borowski p engine
Thrust4.4×105 N
Thrust Power2.7 GW
Engine Mass7,000 kg
(less p containment)
Specific Impulse1,100 sec
Exhaust Velocity10,790 m/s
Fuelantiprotons (p)
Fuel Mass Flow15 μg/sec
(1.5×10-8 kg/s)
PropellantLH2
Propellant
Mass Flow
41 kg/s

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

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

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

The tungsten also acts as the biological shadow shield.

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

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

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

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

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

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

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

### Gas Core

Gas-Core 5k sec
FuelAntiprotons (p)
Fuel Flow Rate2.25×10-8 kg/sec
PropellantLH2
Propellant Flow Rate0.9 kg/sec
Antimatter TargetTungsten
Annihilation Power4.05 GW
Thrust Power1.08 GW
Specific Impulse5,000 sec
Exhaust Velocity49,050 m/s
Thrust44,000 N
Chamber Mass25,000 kg
Magnetic Coil Mass70,000 kg
Total Engine Mass182,000 kg
T/W Ratio2.5×10-2
Specific Power5.9 kW/kg
Gas-Core 1.25k sec
FuelAntiprotons (p)
Fuel Flow Rate1.125×10-8 kg/sec
PropellantLH2
Propellant Flow Rate14.3 kg/sec
Antimatter TargetTungsten
Annihilation Power2.025 GW
Thrust Power0.27 GW
Specific Impulse1,250 sec
Exhaust Velocity12,260 m/s
Thrust44,000 N
Chamber Mass25,000 kg
Magnetic Coil Mass70,000 kg
Total Engine Mass95,000 kg
T/W Ratio4.7×10-2
Specific Power2.8 kW/kg

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

### Plasma Core

AM: Plasma
Water
Exhaust
Velocity
980,000 m/s
Specific
Impulse
99,898 s
Thrust61,000 N
Thrust
Power
29.9 GW
Mass
Flow
0.06 kg/s
T/W0.01
RemassWater
Specific
Power
17 kg/MW
AM: Plasma
Hydrogen
Exhaust
Velocity
7,840,000 m/s
Specific
Impulse
799,185 s
Thrust49,000 N
Thrust
Power
0.2 TW
Mass
Flow
0.01 kg/s
T/W0.01
RemassLiquid Hydrogen
Specific
Power
3 kg/MW
AM: Plasma
Both
Total
Engine Mass
500,000 kg
FuelAntimatter:
antihydrogen
ReactorPlasma Core
Remass
Accel
Thermal Accel:
Reaction Heat
Thrust
Director
Magnetic Nozzle

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

LaPointe
antiproton
magnetically
confined
plasma
Moderate Density
Hydrogen
Density
1016 atoms/cm3
Antiproton
Density
1010 antiprotons/cm3 to
1012 antiprotons/cm3
Normalized
Thrust
7.6×10-7 N⋅s/cm3 to
9.8×10-6 N⋅s/cm3
Exhaust
Velocity
45,000 m/s to
590,000 m/s
Specific
Impulse
4,610 s to
60,000 s
High Density
Hydrogen
Density
1018 atoms/cm3
Antiproton
Density
1012 antiprotons/cm3
Normalized
Thrust
8.1×10-5 N⋅s/cm3
Exhaust
Velocity
49,000 m/s
Specific
Impulse
4,950 s

LaPointe Antiproton Magnetically Confined Plasma Engine

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

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

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

LaPointe studied a range of densities for the hydrogen propellant.

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

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

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

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

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

T = (Tnormalized / ΔT) * Bvol

where

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

The propellant mass flow is:

mDotp = (mp * np * Bvol) / ΔT

where

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

And obviously the antimatter mass flow is:

mDotp = (mp * np * Bvol) / ΔT

where

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

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

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

Key engineering issues:

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

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

### Beam Core

AM: Beam
Exhaust Velocity100,000,000 m/s
Specific Impulse10,193,680 s
Thrust10,000,000 N
Thrust Power500.0 TW
Mass Flow0.10 kg/s
Total Engine Mass10,000 kg
T/W102
FuelAntimatter:
antihydrogen
ReactorAntimatter Catalyzed
RemassReaction
Products
Remass AccelAnnihilation
Thrust DirectorMagnetic Nozzle
Specific Power2.00e-05 kg/MW

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

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

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

### Positron Ablative

Positron Ablative
Exhaust velocity49,000 m/s

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

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

This system is very similar to Antiproton-catalyzed microfission

## Beamed Power

### Laser Thermal

Laser Thermal
Exhaust Velocity40,000 m/s
Specific Impulse4,077 s
Thrust13,000 N
Thrust Power0.3 GW
Mass Flow0.33 kg/s
Total Engine Mass20,000 kg
T/W0.07
Thermal eff.30%
Total eff.30%
FuelExternal
Laser
ReactorCollector Mirror
RemassSeeded Hydrogen
Remass AccelThermal Accel:
Collector Mirror
Thrust DirectorNozzle
Specific Power77 kg/MW

Similar to Solar Moth, but uses a stationary ground or space-station based laser instead of the sun. Basically the propulsion system leaves the power plant at home and relies upon a laser beam instead of an incredibly long extension cord.

As a general rule, the collector mirror of a laser thermal rocket can be much smaller than a comparable solar moth, since the laser beam probably has a higher energy density than natural sunlight.

With the mass of the power plant not actually on the spacecraft, more mass is available for payload. Or the reduced mass makes for a higher mass ratio to increase the spacecraft's delta V. The reduced mass also increases the acceleration. In some science fiction novels, combat "motherships" will have batteries of lasers, used to power hordes of ultra-high acceleration missiles and/or fighter spacecraft.

The drawback include the fact that there is a maximum effective range you can send a worthwhile laser beam from station to spacecraft, and the fact that the spacecraft is at the mercy of whoever is controlling the laser station.

Propellant is hydrogen seeded with alkali metal. As always the reason for seeding is that hydrogen is more or less transparent so the laser beam will mostly pass right through without heating the hydrogen. The seeding make the hydrogen more opaque so the blasted stuff will heat up. Having said that, the Mirror Steamer has an alternate solution.

The equations for delta V and mass ratio are slightly different for a Solar Moth or Laser Thermal rocket engine:

Δv = sqrt((2 * Bp * Bε) / mDot) * ln[R]

R = ev/sqrt((2 * Bp * Bε) / mDot)

where

• Δv = ship's total deltaV capability (m/s)
• R = ship's mass ratio
• Bp = Beam power (watts) of either laser beam or solar energy collected
• = efficiency with which engine converts beam power into exhaust kinetic energy (0.0 to 1.0, currently about 0.3)
• ln[x] = natural logarithm of x, the "ln" key on your calculator
• ex = antilog base e or inverse of natural logarithm of x, the "ex" key on your calculator

### Laser Sail

A Laser Sail is a photon sail beam-powered by a remote laser installation.

This is a Propellant-less Rocket, and thus is not subject to The Tyranny of the Rocket Equation.

As an important point, the practical minimum acceleration for a spacecraft is about 5 milligees. Otherwise it will take years to change orbits. Photon sails can only do up to 3 milligees, but a laser sail can do 5 milligees easily.

### Solar Moth

Solar Moth
Exhaust Velocity9,000 m/s
Specific Impulse917 s
Thrust4,000 N
Thrust Power18.0 MW
Mass Flow0.44 kg/s
Total Engine Mass100 kg
T/W4
Thermal eff.65%
Total eff. (Bε)65%
FuelSolar Photons
ReactorCollector Mirror
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Collector Mirror
Thrust DirectorNozzle
Specific Power6 kg/MW

Solar thermal rocket. 175 meter diameter aluminum coated reflector concentrates solar radiation onto a window chamber hoop boiler, heating and expanding the propellant through a regeneratively-cooled hoop nozzle. The concentrating mirror is one half of a giant inflatable balloon, the other half is transparent (so it has an attractive low mass).

The advantage is that you have power as long as the sun shines and your power plant has zero mass (as far as the spacecraft mass is concerned). The disadvantage is it doesn't work well past the orbit of Mars. The figures in the table are for Earth orbit.

The solar moth might be carried on a spacecraft as an emergency propulsion system, since the engine mass is so miniscule.

The equations for delta V and mass ratio are slightly different for a Solar Moth or Laser Thermal rocket engine:

Δv = sqrt((2 * Bp * Bε) / mDot) * ln[R]

R = ev/sqrt((2 * Bp * Bε) / mDot)

where

• Δv = ship's total deltaV capability (m/s)
• R = ship's mass ratio
• Bp = Beam power (watts) of either laser beam or solar energy collected
• = efficiency with which engine converts beam power into exhaust kinetic energy (0.0 to 1.0)
• ln[x] = natural logarithm of x, the "ln" key on your calculator
• ex = antilog base e or inverse of natural logarithm of x, the "ex" key on your calculator

For the Solar Moth in the data block Bε = 0.65, for the Mirror Steamer Bε = 0.87

Bp = Marea * (☉constant * (1 / (☉dist2)))

where

• Bp = Beam power (watts) of solar energy collected
• Marea = total area of collecting mirrors (m2)
• dist = distance between Sun and spacecraft (Astronomical Units, Earth = 1.0)
• constant = Solar Constant = varies from 1,361 w/m2 at solar minimum and 1,362 w/m2 at solar maximum (w/m2)

1.0 astronomical units is defined as 149,597,870,700 meters.

1 / (☉dist2) is the sunlight energy density. In Earth's orbit, the density is 1.0, at Mars orbit it is 0.44 (44%), at Jupiter orbit it is 0.037, at Neptune orbit it is 0.001, at Mercury orbit it is 6.68.

Noted space artist Nick Stevens has been working on visualizing a Solar Moth.

## Chemical

A barely contained chemical explosive. Noted for very high thrust and very low exhaust velocity. One of the few propulsion systems where the fuel and the propellant are the same thing. There is a list of chemical propellants here

### Solid Rocket

Exhaust Velocity 2,600 m/s 265 s 12,000,000 N x2 24,000,000 N 31.2 GW 9,231 kg/s 1,180,000 kg 2 Chemical Solid:APCP CombustionChamber ReactionProducts Thermal Accel:Reaction Heat Nozzle 38 kg/MW

### Liquid Rocket

#### Methane-Oxygen

Chemical: Methane-Oxygen
Exhaust Velocity3,700 m/s
Specific Impulse377 s

Methane and oxygen are burned resulting in an unremarkable specific impulse of about 377 seconds. However, this is the highest performance of any chemical rocket using fuels that can be stored indefinitely in space. Chemical rockets with superior specific impulse generally use liquid hydrogen, which will eventually leak away by escaping between the the molecules composing your fuel tanks. Liquid methane and liquid oxygen will stay put.

People tend to sneer at chemical rockets because of their abysmal specific impulse. Surprisingly, they are perfectly adequate for missions to Mars or cis-Lunar space provided there is a network of orbital propellant depots suppled by in-situ resource allocation. See Sabatier reaction below.

#### Hydrogen-Fluorine

Chemical: LH2/Fluorine
Exhaust Velocity4,700 m/s
Specific Impulse479 s

#### Hydrogen-Oxygen

Chemical: LH2/LOX
Exhaust Velocity4,400 m/s
Specific Impulse449 s
Space Shuttle SSME x3
Propulsion SystemChemical: LH2/LOX
Exhaust Velocity4,444 m/s
Specific Impulse453 s
Thrust/Engine1,817,400 N
Number Thrustersx3
Thrust5,452,200 N
Thrust Power12.1 GW
Mass Flow1,227 kg/s
Total Engine Mass9,531 kg
T/W58
Specific Power1 kg/MW
NASA space tug
Propulsion SystemChemical: LH2/LOX
Thrust22,400 N
Thrust Power49.3 MW
Mass Flow5 kg/s
Total Engine Mass199,600 kg
T/W0.01
Wet Mass32,000 kg
Dry Mass14,000 kg
Mass Ratio2.29 m/s
ΔV3,637 m/s
Specific Power4,050 kg/MW
Lighter
Propulsion SystemChemical: LH2/LOX
Exhaust Velocity4,410 m/s
Specific Impulse450 s
Thrust140,000 N
Thrust Power0.3 GW
Mass Flow32 kg/s
Wet Mass56,300 kg
Dry Mass25,898 kg
Mass Ratio2.17 m/s
ΔV3,424 m/s
Kuck Mosquito
Propulsion SystemChemical: LH2/LOX
Exhaust Velocity4,400 m/s
Specific Impulse449 s
Thrust220,000 N
Thrust Power0.5 GW
Mass Flow50 kg/s
Wet Mass350,000 kg
Dry Mass100,000 kg
Mass Ratio3.50 m/s
ΔV5,512 m/s

Hydrogen and oxygen are burned resulting in close to the theoretical maximum specific impulse of about 450 seconds. However, liquid hydrogen cannot be stored permanently in any tank composed of matter. The blasted stuff will escape atom by atom between the molecules composing the fuel tanks.

People tend to sneer at chemical rockets because of their abysmal specific impulse. Surprisingly, they are perfectly adequate for missions to Mars or cis-Lunar space provided there is a network of orbital propellant depots suppled by in-situ resource allocation. Even a single depot in Low Earth Orbit supplied from Lunar ice will be a big help.

#### RP-1 - Oxygen

Chemical: RP-1/LOX
Exhaust Velocity3,500 m/s
Specific Impulse357 s
Saturn-V F-1 x1
Exhaust Velocity2,982 m/s
Specific Impulse304 s
Thrust7,740,500 N
Thrust Power11.5 GW
Mass Flow2,596 kg/s
Total Engine Mass9,153 kg
T/W86
FuelChemical:
RP-1/LOX
ReactorCombustion
Chamber
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power1 kg/MW
Saturn-V F-1 x5
Exhaust Velocity3,000 m/s
Specific Impulse306 s
Thrust/Engine7,740,500 N
Number Thrustersx5
Thrust38,702,500 N
Thrust Power58.1 GW
Mass Flow12,901 kg/s
Total Engine Mass45,765 kg
T/W86
FuelChemical Liquid:
RP-1/LOX
Specific Power1 kg/MW

RP-1 is Rocket Propellant-1 or Refined Petroleum-1) is a highly refined form of kerosene outwardly similar to jet fuel, used as rocket fuel. It is not as powerful as liquid hydrogen but it is a whole lot less trouble. Compared to LH2 it is cheaper, stabler at room temperature, non-cryogenic less of an explosive hazard, and denser.

NASA uses it a lot.

#### Hypergolic Fuels

Chemical: UDMH/N204
Exhaust Velocity3,267 m/s
Specific Impulse333 s
Chemical: MMH/N204
Exhaust Velocity3,296 m/s
Specific Impulse336 s
Space Shuttle RCS
Thrust3,870 N
Thrust Power6.0 MW
Mass Flow1 kg/s
Total Engine Mass4 kg
T/W107
FuelChemical:
MMH/N204
Specific Power1 kg/MW
Proton RD-253 x1
Thrust1,830,000 N
Thrust Power2.8 GW
Mass Flow590 kg/s
Total Engine Mass1,260 kg
T/W148
FuelChemical:
UDMH/N204
Specific Power0.44 kg/MW
Proton RD-253 x6
Thrust/Engine1,745,000 N
Number Thrustersx6
Thrust10,470,000 N
Thrust Power16.2 GW
Mass Flow3,377 kg/s
Total Engine Mass7,560 kg
T/W141
FuelChemical:
UDMH/N204
Specific Power0.47 kg/MW

Unsymmetrical dimethylhydrazine (UDMH) + nitrogen tetroxide (N204 or "NTO") and Monomethylhydrazine (MMH) + NTO are very important chemical rocket fuels.

Both are hypergolic, meaning the stuff explodes on contact with each other instead of needing a pilot light or other ignition system as do other chemical fuels. This means one less point of failure and one less maintenance nightmare on your spacecraft. Being hypergolic also prevents large amounts of fuel and oxidizer accumulating in the nozzle, which can cause a hard start or engine catastrophic failure (fancy term for "engine goes ka-blam!"). It is also non-cryogenic, liquid at room temperature and pressure. This means it is a storable liquid propellant, suitable for space missions that last years.

"Ah, what's the catch?" you ask.

The catch is that the mix is hideously corrosive, toxic, and carcinogenic. It is also easily absorbed through the skin. If UDMH escapes into the air it reacts to form dimethylnitrosamine, which is a persistent carcinogen and groundwater pollutant. MMH is only fractionally less bad.

This is the reason for all those technicians wearing hazmat suits at Space Shuttle landings. The Shuttle used MMH/NTO in its reaction control thrusters. Upon landing the techs had to drain the hellish stuff before it leaked and dissoved some innocent bystander.

In the words of Troy Campbell, hypergolic fuels are tanks full of explosive cancer.

### Hybrid Rocket

#### Aluminum-Oxygen

Chemical: Aluminum-Oxygen
Exhaust Velocity2,800 m/s
Specific Impulse285 s
FuelChemical:
Aluminum/LOX
ReactorCombustion
Chamber
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle

Aluminum and oxygen are burned resulting in an unremarkable specific impulse of about 285 seconds. However, this is of great interest to any future lunar colonies. Both aluminum and oxygen are readily available in the lunar regolith, and such a rocket could easily perform lunar liftoff, lunar landing, or departure from a hypothetical L5 colony for Terra (using a lunar swingby trajectory). The low specific impulse is more than made up for by the fact that the fuel does not have to be imported from Terra. It can be used in a hybrid rocket (with solid aluminum burning in liquid oxygen), or using ALICE (which is a slurry of nanoaluminium powder mixed in water then frozen).

Of course the aluminum oxide in lunar regolith has to be split into aluminum and oxygen before you can use it as fuel. But Luna has plenty of solar power. As a general rule, in space, energy is cheap but matter is expensive.

### Metastable

#### Atomic Hydrogen

100% Atomic Hydrogen
Exhaust velocity20,600 m/s
15% Atomic Hydrogen in solid H2
Exhaust velocity7,300 m/s
Single-H/LOX
Exhaust Velocity4,600 m/s
Specific Impulse469 s

Atomic hydrogen is also called free-radical hydrogen or "single-H". The problem is that it instantly wants to recombine. The least unreasonable way of preventing this is to make a solid mass of frozen hydrogen (H2) at liquid helium temperatures which contains 15% single-H by weight.

#### Metallic Hydrogen

Metallic Hydrogen
Specific Impulse1,700 sec
Exhaust Velocity16,700 m/s
Reaction Chamber
Temperature
6,000 K
Density700 kg/m3
Energy of
Recombination
216 MJ/kg

Most of the data here is from Metallic Hydrogen: The Most Powerful Rocket Fuel Yet to Exist by Isaac F. Silvera and John W. Cole.

Hydrogen (H2) subjected to enough pressure to turn it into metal (mH), then contained under such pressure. Release the pressure and out comes all the stored energy that was required to compress it in the first place.

It will require storage that can handle millions of atmospheres worth of pressure. The mass of the storage unit might be enough to negate the advantage of the high exhaust velocity.

Or maybe not. The hope is that somebody might figure out how to compress the stuff into metal, then somehow release the pressure and have it stay metallic. In Properties of Metallic Hydrogen under Pressure the researchers showed that hydrogen would be a metastable metal with a potential barrier of ~1 eV. That is, if the pressure on metallic hydrogen were relaxed, it would still remain in the metallic phase, just as diamond is a metastable phase of carbon. This will make it a powerful rocket fuel, as well as a candidate material for the construction of Thor's Hammer.

Then that spoil-sport E. E. Salpeter wrote in "Evaporation of Cold Metallic Hydrogen" a prediction that quantum tunneling might make the stuff explode with no warning. Since nobody has managed to make metallic hydrogen they cannot test it to find the answer.

Silvera and Cole figure that metallic hydrogen is stable, to use it as rocket fuel you just have to heat it to about 1,000 K and it explodes recombines into hot molecular hydrogen.

Recombination of hydrogen from the metallic state would release a whopping 216 megajoules per kilogram. TNT only releases 4.2 megajoules per kg. Hydrogen/oxygen combustion in the Space Shuttle main engine releases 10 megajoules/kg. This would give metallic hydrogen an astronomical specific impulse (Isp) of 1,700 seconds. The shuttle only had 460 seconds, NERVA had 800, and the pebble bed NTR had 1,000 seconds. Yes, this means metallic hydrogen has more specific impulse than a freaking solid-core nuclear thermal rocket.

Isp of 1,700 seconds is big enough to build a single-stage-to-orbit heavy lift vehicle, which is the holy grail of boosters.

The cherry on top of the sundae is that metallic hydrogen is about ten times more dense (700 kg/m3) than that pesky liquid hydrogen (70.8 kg/m3). The high density is a plus, since liquid hydrogen's annoyingly low density causes all sorts of problems. Metallic hydrogen also probably does not need to be cryogenically cooled, unlike liquid hydrogen. Cryogenic cooling equipment cuts into your payload mass.

The drawback is the metallic hydrogen reaction chamber will reach a blazing temperature of at least 6,000 K. By way of comparison the temperatures in the Space Shuttle main engine combustion chamber can reach 3,570 K, which is about the limit of the state-of-the-art of preventing your engine from evaporating.

It is possible to lower the combustion chamber temperature by injecting cold propellant like water or liquid hydrogen. The good part is you can lower the temperature to 3,570 K so the engine doesn't melt. The bad part is this lowers the specific impulse (nothing comes free in this world). But even with a lowered specific impulse the stuff is still revolutionary.

At 100 atmospheres of pressure in the combustion chamber it will be an Isp of 1,700 sec with a temperature of 7,000 K. At 40 atmospheres the temperature will be 6,700 K, still way to high.

Injecting enough water propellant to bring the temperature down to 3,500 to 3,800 K will lower the Isp to 460 to 540 seconds. Doing the same with liquid hydrogen will lower the Isp to 1,030 to 1,120 seconds.

 Dilutant Isp (s) ChamberTemp (K) Mix Ratio(H2/mH) - H2 H2 H2 H2 H2 H2 H2 H2 H2O H2O H2O H2O 1700 1091? 1120 1089 1058 1029 1022 962 911 538 512 489 467 7000 3925 3800 3700 3600 3500 3673 3448 3240 3800 3700 3600 3500 - 1.50 1.87 2.09 2.33 2.59 2.00 2.50 3.00 10.76 12.22 13.79 15.44

#### Metastable He*

Metastable He*
Exhaust Velocity43,000 m/s
Specific Impulse4,383 s
Thrust64,000 N
Thrust Power1.4 GW
Mass Flow1 kg/s
Total Engine Mass10,000 kg
T/W0.65
FuelMetastable He*
ReactorCombustion
Chamber
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power7 kg/MW

Spin-polarized triplet helium. Two electrons in a helium atom are aligned in a metastable state (one electron each in the 1s and 2s atomic orbitals with both electrons having parallel spins, the so-called "triplet spin state", if you want the details). When it reverts to normal state it releases 0.48 gigjoules per kilogram. Making the stuff is easy. The trouble is that it tends to decay spontaneously, with a lifetime of a mere 2.3 hours. And it will decay even quicker if something bangs on the fuel tank. Or if the ship is jostled by hostile weapons fire. To say the fuel is touchy is putting it mildly. The fuel is stored in a resonant waveguide to magnetically lock the atoms in their metastable state but that doesn't help much. There were some experiments to stablize it with circularly polarized light, but I have not found any results about that.

#### Metastable He IV-A

Metastable He IV-A
Exhaust Velocity21,600 m/s
Specific Impulse2,202 s
Total Engine Mass10,000 kg
FuelMetastable He IV-A
ReactorCombustion
Chamber
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Meta from Saturn Rukh
Exhaust Velocity30,900 m/s
Specific Impulse3,150 s

Meta-helium would be such a worthwhile propulsion system that scientists have been trying real hard to get the stuff to stop decaying after a miserable 2.3 hours. One approach is to see if metastable helium can be formed into a room-temperature solid if bonded with diatomic helium molecules, made from one ground state atom and one excited state atom. This is called diatomic metastable helium. The solid should be stable, and it can be ignited by heating it. The exhaust velocity is about half that of pure He* which is disappointing, but not as disappointing as a dust-mote sized meteorite blowing your ship into atoms.

Theoretically He IV-A would be stable for 8 years, have a density of 0.3 g/cm3, and be a solid with a melting point of 600 K (27° C). The density is a plus, liquid hydrogen's annoying low density causes all sorts of problems.

Dr. Robert Forward in his novel Saturn Rukh suggested bonding 64 metastable helium atoms to a single excited nitrogen atom, forming a stable super-molecule called Meta. Whether or not this is actually possible is anybody's guess. In theory it would have a specific impulse of 3150 seconds.

## Electromagnetic (Plasma)

Electromagnetic ion thrusters use the Lorentz force to move the propellant ions.

### Helicon Double Layer (HDLT)

Exhaust Velocity314,000 m/s
Specific Impulse32,008 s
Thrust20,000 N
Thrust Power3.1 GW
Mass Flow0.06 kg/s
Total Engine Mass1,540,000 kg
T/W1.00e-03
Thermal eff.79%
Total eff.79%
Fuel4GWe input
RemassHelium
Remass AccelElectromagnetic
Acceleration
Specific Power490 kg/MW
HOPE Cargo MPD
Propulsion SystemMPD
Exhaust Velocity78,500 m/s
Specific Impulse8,002 s
Thrust11 N
Number Thrusters2
Thrust Power0.4 MW
Mass Flow1.40e-04 kg/s
Fuel60MWe input
RemassHelium
Wet Mass242,000 kg
Dry Mass182,000 kg
Mass Ratio1.33 m/s
ΔV22,367 m/s
HOPE Tanker MPD
Propulsion SystemMPD
Exhaust Velocity78,500 m/s
Specific Impulse8,002 s
Thrust11 N
Number Thrusters2
Thrust Power0.4 MW
Mass Flow1.40e-04 kg/s
Fuel60MWe input
RemassHelium
Wet Mass244,000 kg
Dry Mass184,000 kg
Mass Ratio1.33 m/s
ΔV22,155 m/s
HOPE Crew MPD
Propulsion SystemMPD
Exhaust Velocity78,500 m/s
Specific Impulse8,002 s
Thrust28 N
Number Thrusters4
Thrust Power1.1 MW
Mass Flow3.57e-04 kg/s
Fuel60MWe input
RemassHelium
Wet Mass262,000 kg
Dry Mass188,000 kg
Mass Ratio1.39 m/s
ΔV26,054 m/s

Magnetoplasmadynamic thruster, a travelling wave plasma accelerator. Propellant is potassium seeded helium.

### Pulsed Inductive (PIT)

Pulsed inductive thruster

Paper here

### VASIMR

VASIMR
VASIMR (high gear)
Exhaust Velocity294,000 m/s
Specific Impulse29,969 s
Thrust40 N
Thrust Power5.9 MW
Mass Flow1.36e-04 kg/s
Total Engine Mass10,000 kg
T/W4.08e-04
Specific Power1,701 kg/MW
VASIMR (med gear)
Exhaust Velocity147,000 m/s
Specific Impulse14,985 s
Thrust80 N
Thrust Power5.9 MW
Mass Flow5.44e-04 kg/s
Total Engine Mass10,000 kg
T/W8.15e-04
Specific Power1,701 kg/MW
VASIMR (low gear)
Exhaust Velocity29,000 m/s
Specific Impulse2,956 s
Thrust400 N
Thrust Power5.8 MW
Mass Flow0.01 kg/s
Total Engine Mass10,000 kg
T/W4.08e-03
Specific Power1,724 kg/MW
All
Thermal eff.60%
Total eff.60%
Fuel19.6MWe input
RemassLiquid Hydrogen
Remass AccelElectromagnetic
Acceleration
Thrust DirectorMagnetic Nozzle

Some classify this as an electromagnetic plasma, some as an electrodeless electrothermal

The variable specific impulse magnetoplasma rocket is a plasma drive with the amusing ability to "shift gears." This means it can trade exhaust velocity for thrust and vice versa. Three "gears" are shown on the table. There are more details here and here.

VASIMR has been suggested for use in a space tug aka Orbital Transfer Vehicle. A VASIMR powered tug could move 34 metric tons from Low Earth Orbit (LEO) to Low Lunar Orbit (LLO) by expending only 8 metric tons of argon propellant. A chemical rocket tug would require 60 metric tons of liquid oxygen - liquid hydrogen propellant. Granted the VASIMR tug would take six month transit time as opposed to the three days for the chemical, but there are always trade offs.

## Electrostatic

Electrostatic ion thrusters use the Coulomb force to move the propellant ions.

What the joke is saying is that electrostatic drives are power hogs. Solar power is relatively lightweight but the energy is so dilute you need huge arrays. Nuclear power can supply megawatts of power, but reactors have a mass measured in tons.

But the joke is on the wag. Turns out there is such a thing as an extension cord long enough, it is called beamed power. This is where the spacecraft has a relatively lightweight power receptor, while back at home is a kilometers-wide solar power station that gathers gigawatts of power and beams it to the spacecraft via microwave beam or laser. The beam becomes the extension cord.

### Electrostatic Propellant

When I was a little boy, the My First Big Book of Outer Space Rocketships type books I was constantly reading usually stated that ion drives would use mercury or cesium as propellant. But most NASA spacecraft are using xenon. What's the story?

Ionization energy represents a large percentage of the energy needed to run ion drives. The ideal propellant is thus easy to ionize and has a high mass/ionization energy ratio. In addition, the propellant should not erode the thruster to any great degree to permit long life; and should not contaminate the vehicle.

Many current designs use xenon gas, as it is easy to ionize, has a reasonably high atomic number, is inert and causes low erosion. However, xenon is globally in short supply and expensive.

Older designs used mercury, but this is toxic and expensive, tended to contaminate the vehicle with the metal and was difficult to feed accurately.

Other propellants, such as bismuth and iodine, show promise, particularly for gridless designs, such as Hall effect thrusters.

Gridded Electrostatic Ion Thrusters typically use xenon.

Hal Effect Thrusters typically use xenon, bismuth and iodine

Field-Emission Electric Propulsion typically use caesium or indium as the propellant due to their high atomic weights, low ionization potentials and low melting points.

Pulsed Inductive Thrusters typically use ammonia gas.

Magnetoplasmadynamic Thrusters typically use hydrogen, argon, ammonia or nitrogen.

If you want the ultimate in in-situ resource utilization, design an ion drive that can use asteroid dust for propellant.

### Colloid

ESTAT: Colloid
Exhaust Velocity43,000 m/s
Specific Impulse4,383 s
Thrust8,000 N
Thrust Power0.2 GW
Mass Flow0.19 kg/s
Total Engine Mass20,000 kg
T/W0.04
Thermal eff.85%
Total eff.85%
Fuel200MWe input
RemassColloid
Remass AccelElectrostatic
Acceleration
Specific Power116 kg/MW

Similar to Ion, but utilizing tiny droplets instead of ions.

### Field-Emission Electric (FEEP)

Field-emission electric propulsion, a type of Colloid thruster.

They typically use caesium or indium as the propellant due to their high atomic weights, low ionization potentials and low melting points.

### Hall Effect (HET)

Hall Effect Thruster

### Ion

Ion
Exhaust Velocity210,000 m/s
Specific Impulse21,407 s
Thrust10,000 N
Thrust Power1.1 GW
Mass Flow0.05 kg/s
Total Engine Mass400,000 kg
T/W3.00e-03
Thermal eff.96%
Total eff.96%
Fuel800MWe input
RemassArgon
Remass AccelElectrostatic
Acceleration
Specific Power381 kg/MW
DAWN mission NSTAR
Propulsion SystemIon
Exhaust Velocity30,411 m/s
Specific Impulse3,100 s
Thrust9.00e-05 N
Thrust Power1.4 W
Mass Flow2.96e-09 kg/s
Total Engine Mass26 kg
T/W3.60e-07
FuelSolar Photons
ReactorPhotovoltaic array
RemassXenon
Remass AccelElectrostatic
Acceleration
Wet Mass1,210 kg
Dry Mass785 kg
Mass Ratio1.54 m/s
ΔV13,159 m/s
Specific Power1.86e+07 kg/MW
Umbrella Ship
Propulsion SystemIon
Exhaust Velocity80,442 m/s
Specific Impulse8,200 s
Thrust490 N
Thrust Power19.7 MW
Mass Flow0.01 kg/s
FuelFission:
Uranium 235
ReactorNuclear Power
Reactor (electric)
RemassCesium
Remass AccelElectrostatic
Acceleration
Wet Mass660,000 kg
Dry Mass328,000 kg
Mass Ratio2.01 m/s
ΔV56,247 m/s

Gridded Electrostatic Ion Thruster. Potassium seeded argon is ionized and the ions are accelerated electrostatically by electrodes. Other propellants can be used, such as cesium and buckyballs. Though it has admirably high exhaust velocity, there are theoretical limits that ensure all Ion drives are low thrust.

It also shares the same problem as the other electrically powered low-thrust drives. In the words of a NASA engineer the problem is "we can't make an extension cord long enough." That is, electrical power plants are weighty enough to make the low thrust an even larger liability. A high powered ion drive will generally be powered by a nuclear reactor, Nuclear Electric Propulsion (NEP). Low powered ion drives can get by with solar power arrays, all ion drive space probes that exist in the real world use that system. Researchers are looking into beamed power systems, where the ion drive on the spaceship is energized by a laser beam from a remote space station.

If you are interested in the technical details about why ion drives are low thrust, read on.

### ( IBS Agamemnon )

IBS Agamemnon
Propulsion SystemIon
Exhaust Velocity220,000 m/s
Specific Impulse22,426 s
Thrust10,000,000 N
Thrust Power1.1 TW
Mass Flow45 kg/s
FuelDeuterium-Deuterium
Fusion
ReactorFusion Power
Reactor(electric)
Remass AccelElectrostatic
Acceleration
Wet Mass100,000,000 kg
Dry Mass28,000,000 kg
Mass Ratio3.57 m/s
ΔV280,052 m/s
Ship Mass8,000,000 kg
Cargo Mass20,000,000 kg
Length400 m
Length spin arm100 m
Cargo Tug Slingshot
Propulsion SystemIon
Exhaust Velocity280,000 m/s
Specific Impulse28,542 s
Thrust5,460,000 N
Thrust Power764.4 GW
Mass Flow20 kg/s
FuelDeuterium-Deuterium
Fusion
ReactorFusion Power
Reactor(electric)
Remass AccelElectrostatic
Acceleration
Wet Mass512,600,000 kg
Dry Mass501,600,000 kg
Mass Ratio1.02 m/s
ΔV6,074 m/s

Fictional Interplanetary BoostShip Agamemnon from Jerry Pournelle's short story "Tinker". This fictional ship is a species of Ion drive utilizing cadmium and powered by deuterium fusion. Looking at its performance I suspect that in reality no Ion drive could have such a high thrust. The back of my envelope says that you'd need one thousand ultimate Ion drives to get this much thrust.

## Electrothermal

### ArcJet

ArcJet
Exhaust Velocity20,000 m/s
Specific Impulse2,039 s
Thrust2 N
Thrust Power20.0 kW
Mass Flow1.00e-04 kg/s
Fuel100kWe input
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Arc Heater
Thrust DirectorNozzle

Hydrogen propellant is heated by an electrical arc.

### Microwave Electrothermal

Microwave Electrothermal Thruster

### Resistojet

Resistojet
Exhaust Velocity2,900 m/s
Specific Impulse296 s
Thrust1 N
Thrust Power0.7 kW
Mass Flow2.00e-04 kg/s
Thermal eff.80%
Total eff.80%
Fuel100kWe input
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Resistance Heater
Thrust DirectorNozzle

In a resistojet, propellant flows over a resistance-wire heating element (much like a space heater or toaster) then the heated propellant escapes out the exhaust nozzle. They are mostly used as attitude jets on satellites, and in situations where energy is more plentiful than mass.

## Fusion

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

• The exhaust velocity/specific impulse is attractively high

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

Drawbacks include:

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

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

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

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

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

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

### Fusion Fuels

For more details about fusion fuels, go here.

#### Deuterium-Tritium

Exhaust Velocity 22,000 m/s 2,243 s 108,000 N 1.2 GW 5 kg/s 10,000 kg 1 Deuterium-TritiumFusion 8 kg/MW

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

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

#### Hydrogen-Boron

H-B Fusion
Exhaust Velocity980,000 m/s
Specific Impulse99,898 s
Thrust61,000 N
Thrust Power29.9 GW
Mass Flow0.06 kg/s
Total Engine Mass300,000 kg
T/W0.02
FuelHydrogen-Boron
Fusion
Specific Power10 kg/MW

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

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

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

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

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

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

#### Helium3-Deuterium

He3-D Fusion
Exhaust Velocity7,840,000 m/s
Specific Impulse799,185 s
Thrust49,000 N
Thrust Power0.2 TW
Mass Flow0.01 kg/s
Total Engine Mass1,200,000 kg
T/W4.00e-03
FuelHelium3-Deuterium
Fusion
Specific Power6 kg/MW

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

### Confinement

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

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

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

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

Distance

DAB (A.U.)
Mass
Ratio
RM
Initial
Mass
Mi (mT)
Propellant
Mass
Mp (mT)
Mass
ML/Mi (%)
Travel
Time
τAB (days)
Initial
Acceleration
ai (10-3 g0)
Mars0.5241.7322,1359029.433.0~2.9
Ceres1.7672.4973,0791,8466.569.42.0
Jupiter4.2033.5904,4273,1944.5120.0~1.4
Sample Tokamak Fusion Rocket
Round-trip trajectory
Mission

Mass
Ratio

RM
Propellant
Mass
MpA→B
(mT)
Propellant
Mass
MpB→A
(mT)
Propellant
Mass
MpA→A
(mT)
Initial
Mass
Mi
(mT)
Travel
Time
τAB
(days)
Travel
Time
τBA
(days)
Travel
Time
τRT
(days)
Mars2.6641,1499022,0513,28443.233.977.1
Ceres4.6672,6751,8464,5215,754100.569.4169.9
Jupiter7.7835,1693,1948,3639,596194.3120.0314.3
Sample Inertial Confinement Fusion Rocket
Round-trip continuous-burn constant-Isp trajectory
Mission

Distance

DAB (A.U.)
Mass
Ratio

RM
Initial
Mass
Mi
(mT)
Propellant
Mass
MpA→A
(mT)
Mass
ML/Mi (%)
Travel
Time
τAB
(days)
Travel
Time
τRT
(days)
Mars0.5241.104757.371.326.427.755.0
Ceres1.7671.196820.5134.524.453.1103.7
Jupiter4.2031.309898212.022.384.6163.6
Saturn8.5391.453997311.020.1125.5239.8
Uranus18.1821.6891,159473.017.3194.1364.7
Neptune29.0581.9011,304618.015.3257.3476.9
Pluto38.5182.0631,415729.014.1306.6562.7

The above tables were calculated with the following equations:

Wf = Mf * g0

MB = Mf + MpB→A

1 / α = Mi / MB

1 / β = MB / Mf

Pf = Mp / Mi

RM = 1 / (α * β) (two way)

RM = 1 / β (one way)

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

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

τRT = τAB + τBA (equation 12a)

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

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

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

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

where:

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

#### Inertial Confinement

Inertial Confinement Fusion is in the Pulse section.

#### Magnetic Confinement

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

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

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

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

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

Also known as "Open-field magnetic confinement".

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

##### Toroidal Fusion

Also known as "Closed-field magnetic confinement".

### Fusion Engines

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

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

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

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

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

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

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

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

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

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

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

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

#### Pure Fusion Engines

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

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

The disadvantage it the absurdly small thrust.

To calculate the exhaust velocity of a Pure Fusion Rocket:

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

where

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

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

Ve = sqrt(2 * Ep)

where

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

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

#### Afterburner Fusion Engines

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

This is based on information from physicist Luke Campbell.

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

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

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

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

eq.2     mDot = F2 / (1.7 * Ptherm)

eq.3     Ptherm = F2 / (1.7 * mDot)

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

eq.5     Ve = F / mDot

eq.6     mDot = F / Ve

where:

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

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

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

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

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

#### Dual-Mode Fusion Engines

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

This is based on information from physicist Luke Campbell.

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

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

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

### Nuclear Magnetic Spin Alignment

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

### ( STARFIRE Fusion Afterburner )

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

### ( AV:T Fusion )

AV:T Fusion
Cruise mode
Exhaust Velocity832,928 m/s
Specific Impulse84,906 s
Thrust245,250 N
Thrust Power0.1 TW
Mass Flow0.29 kg/s
FuelHelium3-Deuterium
Fusion
Combat mode
Exhaust Velocity104,116 m/s
Specific Impulse10,613 s
Thrust48,828,125 N
Thrust Power2.5 TW
Mass Flow469 kg/s
FuelHelium3-Deuterium
Fusion

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

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

### ( THS Fusion Pulse )

Fusion Pulse low gear
Exhaust Velocity150,000 m/s
Specific Impulse15,291 s
Thrust80,000 N
Mass Flow0.53 kg/s
T/W2
Fusion Pulse high gear
Exhaust Velocity300,000 m/s
Specific Impulse30,581 s
Thrust40,000 N
Mass Flow0.13 kg/s
T/W1
Both
Thrust Power6.0 GW
Total Engine Mass4,000 kg
FuelHelium3-Deuterium
Fusion
Specific Power1 kg/MW

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

### ( Epstein Drive )

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

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

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

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

Thus ends the canon knowledge.

### My Analysis

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

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

What does those assumptions give us?

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

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

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

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

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

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

### Scott Manley's Analysis

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

### Monstah's Analysis

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

### Erin Schmidt's Analysis

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

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

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

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

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

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

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

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

## Nuclear Thermal

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

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

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

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

### Exhaust Velocity Limits on Nuclear Thermal Rockets

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

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

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

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

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

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

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

### Solid Core

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

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

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

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).

#### NERVA

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.

NERVA
Thrust Power0.198-0.065 GW
Exhaust velocitySee Table
Thrust49,000 n
Engine mass10 tonne
T/W >1.0no
NERVA (H2)
Exhaust Velocity8,093 m/s
Specific Impulse825 s
Thrust49,000 N
Thrust Power0.2 GW
Mass Flow6 kg/s
Total Engine Mass10,000 kg
T/W0.50
FuelFission:
Uranium 235
ReactorSolid Core
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power50 kg/MW
Resuable Nuclear Shuttle [+]
Propulsion SystemNERVA
Exhaust Velocity8,000 m/s
Specific Impulse815 s
Thrust344,000 N
Thrust Power1.4 GW
Mass Flow43 kg/s
Wet Mass170,000 kg
Dry Mass30,000 kg
Mass Ratio5.67 m/s
ΔV13,877 m/s
Widmer Mars Mission [+]
Propulsion SystemNERVA
Exhaust Velocity8,000 m/s
Specific Impulse815 s
Thrust580,000 N
Thrust Power2.3 GW
Mass Flow72 kg/s
Wet Mass400,000 kg
Dry Mass150,000 kg
Mass Ratio2.67 m/s
ΔV7,847 m/s
HELIOS 2nd Stage [+]
Propulsion SystemNTR Solid
Exhaust Velocity7,800 m/s
Specific Impulse795 s
Thrust981,000 N
Thrust Power3.8 GW
Mass Flow126 kg/s
Wet Mass100,000 kg
Dry Mass6,800 kg
Mass Ratio14.71 m/s
ΔV20,968 m/s
Atomic V-2 [+]
Propulsion SystemNTR Solid
Exhaust Velocity8,980 m/s
Specific Impulse915 s
Thrust1,050,000 N
Thrust Power4.7 GW
Mass Flow117 kg/s
Total Engine Mass4,200 kg
T/W25
Wet Mass42,000 kg
Dry Mass17,000 kg
Mass Ratio2.47 m/s
ΔV8,122 m/s
Specific Power1 kg/MW

#### Pewee-class

Pewee-class Engine
Exhaust Velocity9,200 m/s
Specific Impulse940 s
Thrust111,200 N
(25 klbf)
Thrust Power512 MWt
Mass Flow12.5 kg/s
Total Engine Mass3,240 kg
T/W3.5
FuelFission:
Uranium 235
Max Fuel Temp2940 K
Fuel Element
Length
1.32 m
U-235 Mass36.8 kg
Chamber Pressure1000 psi
ReactorSolid Core
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Engine Length
(inc. skirt ext)
7.01 m
Nozzle Skirt
Extension
2.16 m
Nozzle Exit Dia1.87 m
Specific Power6.3 kg/MW
Longest Single
Burn
44.5 min
Total Burn
Duration
79.2 min
Num Burns4

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

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

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

#### SNRE-class

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

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

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

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

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

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

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

Criticality-limited Engine
Exhaust Velocity8,770 m/s
Specific Impulse894 s
Thrust33,000 N
(7.4 klbf)
Thrust Power145 MWt
Mass Flow3.8 kg/s
Total Engine Mass1,770 kg
T/W1.9
FuelFission:
Uranium 235
Max Enrichment93% U-235 wt
Num Fuel Elements260
Num Tie-tube
Elements
251
Max Fuel Temp2736 K
U-235 Mass27.5 kg
ReactorSolid Core
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power12.2 kg/MW
Engine Length6.19 m
Fuel Length
Pressure Vessel
Diameter
0.877 m
Nozzle Exit
Diameter
1.32 m
Longest Single
Burn
22 min
Total Burn
Duration
29.5 min
Num Burns2
SNRE-class Engine
Exhaust Velocity8,829 m/s
Specific Impulse900 s
Thrust73,000 N
(16.7 klbf)
Thrust Power367 MWt
Mass Flow8.4 kg/s
Total Engine Mass2,400 kg
T/W3.06
FuelFission:
Uranium 235
Max Enrichment93% U-235 wt
Num Fuel Elements564
Num Tie-tube
Elements
241
Max Fuel Temp2,726 K
U-235 Mass59.6 kg
ReactorSolid Core
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power6.5 kg/MW
Engine Length4.46 m
Fuel Length0.89 m
Pressure Vessel
Diameter
0.98 m
Nozzle Exit
Diameter
2.26 m
Longest Single
Burn
21.4 min
Total Burn
Duration
55 min
Num Burns5

#### NERVA Derivative

NERVA Deriv
Exhaust Velocity8,085 m/s
Specific Impulse824 s
Thrust334,061 N
Thrust Power1.4 GW
Mass Flow41 kg/s
Total Engine Mass10,100 kg
T/W3
FuelFission:
Uranium 235
ReactorSolid Core
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power7 kg/MW

#### DUMBO

General Dumbo
Thrust Power14.0-4.6 GW
Exhaust velocitySee Table
Thrust3,500,000 n
Engine mass5 tonne
T/W >1.0yes
Dumbo (H2)
Exhaust Velocity8,093 m/s
Specific Impulse825 s
Thrust3,500,000 N
Thrust Power14.2 GW
Mass Flow432 kg/s
Total Engine Mass5,000 kg
T/W71
FuelFission:
Uranium 235
ReactorSolid Core
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Dumbo Model A
Engine mass0.7 tonne
Thrust400,000 n
Propellant mass flow52 kg/sec
Exhaust velocity7,700 m/sec
Engine Height0.6 m
Engine Volume0.2 m3
T/W58
Dumbo Model B
Engine mass2.8 tonne
Thrust3,560,000 n
Propellant mass flow460 kg/sec
Exhaust velocity7,700 m/sec
Engine Height0.6 m
Engine Volume1.8 m3
T/W130
Dumbo Model C
Engine mass2.1 tonne
Thrust400,000 n
Propellant mass flow48 kg/sec
Exhaust velocity8,300 m/sec
Engine Height0.6 m
Engine Volume0.3 m3
T/W20

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

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

#### Pebble Bed

Pebble Bed
Exhaust Velocity9,530 m/s
Specific Impulse971 s
Thrust333,617 N
Thrust Power1.6 GW
Mass Flow35 kg/s
Total Engine Mass1,700 kg
T/W20
FuelFission:
Uranium 235
ReactorSolid Core
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power1 kg/MW

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

#### Cermet

Cermet NERVA
Exhaust Velocity9,120 m/s
Specific Impulse930 s
Thrust445,267 N
Thrust Power2.0 GW
Mass Flow49 kg/s
Total Engine Mass9,000 kg
T/W5
FuelFission:
Uranium 235
ReactorSolid Core
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power4 kg/MW

#### Pulsed Solid-core NTR

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

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

Pulse mode can be used two ways:

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

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

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

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

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

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

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

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

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

How Does It Work?

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

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

Amplification Factor

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

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

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

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

N ≅ 6×10-3 * [ ΔT/t ]

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

Direct Thrust Amplification Details

Thrust power is:

Fp = (F * Ve ) / 2

Thrust is:

F = mDot * Ve

Specific Impulse is:

Isp = Ve / g0

where:

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]

where:

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!

#### Project Timberwind

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

NTR Comparison
NERVATimberwind
45
Engine Mass6,803 kg1,500 kg
Thrust (Vac)333.6 kN392.8 kN
Specific Impulse850 s1,000 s
Burn Time1,200 s449 s
T/W530 !!!

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

Timberwind
45
Timberwind
75
Timberwind
250
Diameter4.25 m2.03 m8.70 m
Thrust (Vac)392.8 kN735.5 kN2,451.6 kN
Specific Impulse1,000 s1,000 s1,000 s
Engine Mass1,500 kg2,500 kg8,300 kg
T/W303030
Burn Time449 s357 s493 s

#### Russian Twisted Ribbon

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

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

Twisted Ribbon Engines
RD-0140NPPS
Thrust (vac) (kN)35.2868
PropellantH2 + HexaneH2
Propellant Mass Flow (kg/s)~4~7.1
Specific Impulse (vac) (s)~900~920
Core outlet temparture (K)3,0002,800 to 2,900
Chamber Pressure (105 Pa)7060
U235 enrichment (%)9090
Fuel Composition(U,Nb,Zn)CU-Zr-C-N
Fuel Element FormTwisted ribbonTwisted ribbon
Generated electrical power (kW)N/A50
Working fluid for power loop
(% by mass)
N/A93% Xe + 7% He
Max temp for power loop (K)N/A1,500
Max press for power loop (105 Pa)N/A9
Working fluid flow rate (kg/s)N/A1.2
Thermal power - propulsion mode (MW)196340
Thermal power - power mode (MW)N/A0.098
Core length (mm)800700
Core diameter (mm)500515
Engine length (mm)3,700No Data
Engine diameter (mm)1,200No Data
Mass (kg)2,000*1,800**

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

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

#### Low Pressure NTR

Engine Mass 835 kg 49,000 newtons 6.0 1,210 sec 9,800 newtons 1.2 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.

LPNTR advantage IMEO for Mars mission
T/WMission
EngineIspEngineEngine
+ shield
Ref
IMEO
500 KM
Earth Orbit
IMEO
Ref NERVA85042.68841400
LPNTR
3200 K
105062.2603814
LPNTR
3600 K
121062.2485611
LPNTR
3600 K
Dual mode
1210
(1350)
6
(1.25)
2.2
(0.44)
440534

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

LANTR
LANTR NERVA mode
Exhaust Velocity9,221 m/s
Specific Impulse940 s
Thrust67,000 N
Thrust Power0.3 GW
Mass Flow7 kg/s
RemassLiquid Hydrogen
LANTR LOX mode
Exhaust Velocity6,347 m/s
Specific Impulse647 s
Thrust184,000 N
Thrust Power0.6 GW
Mass Flow29 kg/s
RemassHydrogen + Oxygen
LANTR Both
FuelFission:
Uranium 235
ReactorSolid Core
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
SpecialLow-High Gear
Nuclear DC-X NERVA
High Gear
Exhaust Velocity9,810 m/s
Specific Impulse1,000 s
Thrust/Engine1,112,000 N
Thrust5,560,000 N
Thrust Power27.3 GW
Mass Flow567 kg/s
T/W3
RemassHydrogen
Specific Power7 kg/MW
Low Gear
Exhaust Velocity5,900 m/s
Specific Impulse601 s
Thrust/Engine3,336,000 N
Thrust16,680,000 N
Thrust Power49.2 GW
Mass Flow2,827 kg/s
T/W9
RemassWater
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power4 kg/MW
Both
Number Thrustersx5
Total Engine Mass199,600 kg
FuelFission:
Uranium 235
SpecialLow-High Gear
Wet Mass460,000 kg

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

#### Bi-Modal NTR

Bimodal NTR Solid (NASA)
Propulsion SystemNTR Solid Bimodal
Exhaust Velocity8,980 m/s
Specific Impulse915 s
Thrust/Engine66,667 N
Number Thrustersx3
Thrust200,000 N
Thrust Power0.9 GW
Mass Flow22 kg/s
Total Engine Mass6,672 kg
T/W3
FuelFission:
Uranium 235
ReactorSolid Core
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
SpecialBimodal
Wet Mass80,000 kg
Dry Mass26,830 kg
Mass Ratio2.98 m/s
ΔV9,811 m/s
Specific Power7 kg/MW

A useful refinement is the Bimodal NTR.

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

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

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

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

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

Pretty ingenious, eh?

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.

#### MITEE

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

##### Basic
Basic MITEE
Exhaust Velocity9,810 m/s
Specific Impulse1,000 s
Thrust14,000 N
Thrust Power68.7 MW
Mass Flow1 kg/s
Total Engine Mass200 kg
T/W7
FuelFission:
Uranium 235
ReactorSolid Core
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power3 kg/MW

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

##### Monatomic H
Monatomic-H MITEE
Exhaust Velocity12,750 m/s
Specific Impulse1,300 s
Thrust2,350 N
Thrust Power15.0 MW
Mass Flow0.18 kg/s
Total Engine Mass200 kg
T/W1
FuelFission:
Uranium 235
ReactorSolid Core
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power13 kg/MW

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

##### Hybrid
HybridMITEE
Exhaust Velocity17,660 m/s
Specific Impulse1,800 s
Thrust1,700 N
Thrust Power15.0 MW
Mass Flow0.10 kg/s
Total Engine Mass10,000 kg
T/W0.02
FuelFission:
Uranium 235
ReactorSolid Core
RemassSingle-H
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power666 kg/MW

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

### Liquid Core

Liquid Core 1
Exhaust Velocity16,000 m/s
Specific Impulse1,631 s
Thrust7,000,000 N
Thrust Power56.0 GW
Mass Flow438 kg/s
Total Engine Mass70,000 kg
T/W10
FuelFission:
Uranium 235
ReactorLiquid Core
RemassWater
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power1 kg/MW
Liquid Core 2
Exhaust velocity14,700 to 25,500 m/s

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

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

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

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

#### LARS

LARS
Exhaust Velocity19,620 m/s
Specific Impulse2,000 s
Thrust20,000 N
Thrust Power0.2 GW
Mass Flow1 kg/s
Total Engine Mass1,000 kg
T/W2
FuelFission:
Uranium 235
ReactorLiquid Core
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power5 kg/MW
Luna
Propulsion SystemLARS
Exhaust Velocity10,300 m/s
Specific Impulse1,050 s
Thrust11,000,000 N
Thrust Power56.6 GW
Mass Flow1,068 kg/s
Total Engine Mass9,000 kg
T/W125
FuelFission:
Uranium 235
RemassWater
Thrust DirectorNozzle
Wet Mass226,000 kg
Dry Mass45,000 kg
Mass Ratio5.02 m/s
ΔV16,623 m/s

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

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

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

### Droplet Core

Droplet Core
Reactor inner diameter1 m
Reactor outer diameter2 m
Reactor inner length3 m
Reactor outer length4 m
Engine length13 m
T/W
5.0
T/W
1.6
Engine mass
6,800 kg
Engine mass
21,200 kg
Engine pressure500 atm
Internal temp6,000K
Isp2,000 sec
Exhaust velocity19,600 m/s
Engine power1,500 MWth
Thrust333,000 N

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

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

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

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

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

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

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

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

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

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

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

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

### Vapor Core

Vapor Core
Thrust Power1.6 GW
Exhaust velocity9,800 to
11,800 m/s
Thrust330,000 n
Propellant mass flow30 kg/sec
Reactor thermal power1,400 to
1,800 MW
Total engine mass6.83 tonne
Fuel element mass total1.35 tonne
Forward reflector mass0.60 tonne
Aft reflector mass0.51 tonne
Total reactor mass5.83 tonne
Misc. engine
component mass
0.9 tonne
T/W5
FuelFission:
Uranium Hexafluoride
ReactorVapor Core
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power4 kg/MW

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

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

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

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

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

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

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

### Gas Core

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

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

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

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

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

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

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

#### Closed Cycle

Gaseous Core NTR closed 1
Exhaust Velocity20,405 m/s
Specific Impulse2,080 s
Thrust445,000 N
Thrust Power4.5 GW
Mass Flow22 kg/s
Total Engine Mass56,800 kg
T/W0.80
FuelFission:
Uranium Hexafluoride
ReactorGas Core
Closed-Cycle
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power13 kg/MW
Gaseous Core NTR closed 2
Thrust Power0.6 to 231 GW
Exhaust velocity10,800 to 31,400 m/s
Thrust117,700 to 14,700,000 n
Engine mass30 to 300 tonne
Engine T/W0.4 to 5.0
Operating Pressure400 to 1600 atm
NASA report nuclear lightbulb
Thrust Power3.7 GW
Engine Power4.6 GW
Exhaust velocity18,300 m/s
Thrust409,000 n
Engine mass32 tonne
Engine T/W1.3
Operating Pressure500 atm
Propellant mass flow22.3 kg/s
Liberty Ship
Propulsion SystemNuclear Lightbulb
Exhaust Velocity30,000 m/s
Specific Impulse3,058 s
Thrust/Engine5,340,000 N
Number Thrustersx7
Thrust37,380,000 N
Thrust Power560.7 GW
Mass Flow1,246 kg/s
Total Engine Mass378,000 kg
T/W10
FuelFission:
Uranium Hexafluoride
Wet Mass2,700,000 kg
Dry Mass1,600,000 kg
Mass Ratio1.69 m/s
ΔV15,697 m/s
Specific Power0.67 kg/MW

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

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.

##### NASA Report

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

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

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

##### UAC Report

The information comes from a series United Aircraft Corporation reports written mostly by Thomas L. Latham. There are more reports than the ones I've used.

The reference design had seven cells with six surrounding the center cell. The entire engine was sized to fit into the Space Shuttle cargo bay. It was also sized at 4.6 gigawatts, 409,000 Newtons, and a specific impulse of 1,860 seconds in order to avoid the need for external heat radiators. At this level no radiators are required for the moderator or pressure vessel, open-cycle cooling will suffice. Above a specific impulsle of 1,860 seconds radiators will be needed or the engine will melt.

If the specific impulse is above 2,500 seconds the nozzle throats will require their own cooling system.

The hydrogen propellant is seeded with tiny tungsten particles due to the unfortunate fact that hydrogen is transparent to the frequencies emitted by the nuclear reaction. Otherwise the chamber walls would be heated instead of the propellant, which is the exact opposite of what we want. The fissioning U235 or U233 fuel also emits ultraviolet light that degrades the transparency of the enclosing quartz "lightbulb." The researchers were experimenting with seeding the uranium with something that would turn the UV into infrared in order to protect the quartz. Happily the ionizing radiation does expose the degraded quartz to a radiation damage annealing effect that restores transparency to some extent.

The fuel is in the form of Uranium Hexafloride.

The average dose rate in the filament-wound fiberglas pressure vessel was calculated to be 0.17 mrad/sec. This would allow about six full-power runs of 1000-sec duration (about 17 minutes) before the total dose became 1000 mrad, the estimated allowable dosage before degradation of the laminate strength commences.

PRIMARY HYDROGEN PROPELLANT CIRCUIT
This is basically the propellant, passing from the propellant tanks to be heated by the nuclear light bulbs, and then rushing through the exhaust nozzles to provide thrust. Along the way it provides some cooling for various items.
Starting at the tank, the primary hydrogen pump sends it through a H2-H2 heat exchanger for preheating (and providing additional heat rejection for the Secondary Hydrogen Circuit). It passes through a H2-Ne heat exchanger to cool off the neon gas in the Neon And Fuel Circuit. It passes through the Fuel And Neon Separator. A turbine then sends it through the Solid Moderator and End-Wall Liners. Somewhere along the line it is seeded with tungsten microparticles so the hydrogen will be heated by the nuclear light bulbs.
Finally it experiences extreme Direct Heating from the nuclear light bulbs, and exits through the exhaust nozzles.
SECONDARY HYDROGEN CIRCUIT
Basically the coolant system. It runs cooling hydrogen over the pressure vessel, nozzles, flow divider, tie rods, liner tubes, and the transparent walls of the quartz light bulbs (during shutdown it also cools the Fuel And Neon Separator).
The now-hot hydrogen passes through a H2-H2 heat exchanger to give the heat to the space radiator. The lukewarm hydrogen passes through a second H2-H2 heat exchanger to cool down further and preheat the propellant hydrogen.
NEON AND FUEL CIRCUIT
The Neon Make-Up supply keeps the neon pressure in the circuit at the required level. The uranium-235 Fuel Make-Up keeps the amount of fuel droplets in the circuit at the required level. Both are fed into the Fuel Cavity in the interior of the quartz light bulbs to create the furious nuclear reaction (unless engine shut-down is in progress, then the Fuel Control Valve closes to shut off the uranium). The reaction provides the direct heating to the Primary Hydrogen Propellant Circuit. Some of the neon goes through the Cavity Bypass Flow.
Only a fraction of the uranium undergoes fission. So the neon/uranium that comes out of the Fuel Cavity is sent through the Neon and Fuel Separator to strain the uranium out of the neon gas. The neon is cooled which makes the uranium gas condense into liquid droplets. The two are separated by a centrifuge. The neon is cooled further by the H2-Ne heat exchanger.
The neon goes to the Neon Pump, the uranium goes to the Fuel Pump and the cycle begins anew.

For additional details see Ref. 5 (Nuclear studies of the nuclear light bulb rocket engine).

Neon supply is the Neon Make-Up supply, keeping the neon pressure in the circuit at the required level. It is fed into the Fuel Cavity (Unit Cavity) tangentally just inside the quartz light bulb Transparent Wall. This creates the neon-uranium vortex.

The Fuel distillation canister is the Fuel Make-Up. It is fed by the Fuel Pump into the fuel injection duct, introducing it into the Fuel Cavity (Unit Cavity). This creates the furious nuclear reaction inside the quartz light bulb, providing the direct heating to the Primary Hydrogen Propellant Circuit.

The mixture of hot neon, unburnt gaseous uranium fuel, and fission products exits the Fuel Cavity via the Exhaust Duct (about two meters long). Not shown is how cool neon is introduced into the entire length of the exhaust duct to [1] cool the exhaust from 6550 K to 1500 K, [2] prevent the exhaust from severely damaging the exhaust duct, [3] condense the gaseous uranium into liquid uranium droplets, and [4] ensuring that the uranium droplets condense inside the neon gas, instead of on the walls of the exhaust duct causing a nuclear reaction.

The 1500 K neon-uranium droplet flow is sent to the Neon and Fuel Separator (Separator) where the two are isolated by a centrifuge. The neon is cooled by the H2-Ne heat exchanger and goes to the Neon Pump. The uranium fuel goes to the Fuel Pump. Alternatively the uranium is distilled to separate out the silicon seeding and the uranium is deposited in the fuel distillation canister.

Values for weight flow rates, temperature, and volume flow rates are indicated at various stations in the system.

In the Neon and Fuel Separator, the seven exhaust duct inlet pipes from the seven nuclear light bulb unit cavities enter from the left. They enter two inlet plenums: four inlet pipes on the top plenum and three on the bottom. Each plenum has an injection slot delivering the gas mix into the separator cavity, with a velocity of 500 m/s at a steep angle designed to spin the gas. The spin centrifugally separates the uranium from the neon, at about 100,000 g's. The uranium is harvested by uranium collector tubes on the separator wall, while the neon is harvested by an outlet pipe on the separator's long axis. The separator cavity and uranium collector tubes have to be maintained at or above 1,500 K, or the uranium will condense on them. This will not only clog the thing up, but if enough uranium plates out it will accumulate a critical mass with regrettable results.

#### Shutdown

1. Close Fuel Injection Control Valve (turn off the uranium)
2. Begin Linear Decrease in Propellant Flow Rate (propellant flow past light bulbs to exhaust nozzles)
3. Begin Linear Increase in Radiator Flow Rate (flow from coolant heat exchanger to radiator)
4. Maintain Secondary Circuit (flow of hydrogen coolant) and Cavity Neon Flow (buffer gas flow inside the quartz light bulbs) at Full Power Value.

Once the engine shut-down sequence is initiated, it takes six seconds for the power level to drop to zero. It only takes 0.8 seconds for power level to drop to 0.01 of full power, during which time the contained uranium fuel drops from the steady-state level of 13.65 kg down to 11.5 kg.

#### Open Cycle

The open-cycle gas core engine has a radioactive exhaust, there is no getting around it. So the first thing you have to do is estimate the radiation hazard and ensure the crew has adequate radiation shielding.

The second thing to do is find a design that does not wastefully allow expensive un-burnt uranium to escape out the tailpipe. Again the general rule is that as long as the hydrogen mass expelled is 25 to 50 times a high as the uranium mass expelled the uranium losses are within acceptable limits. The various open-cycle designs use different strategies to make the hydrogen to uranium exhaust flow ratio as high as possible.

##### General Open Cycle
Open Cycle
Propulsion SystemGas Core NTR
Exhaust Velocity35,000 m/s
Specific Impulse3,568 s
Thrust3,500,000 N
Thrust Power61.2 GW
Mass Flow100 kg/s
Total Engine Mass200,000 kg
FuelFission:
Uranium Hexafluoride
ReactorGas Core
Open-Cycle
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power3 kg/MW
Engine mass30-200 tonne
T/W 11.9 to 1.8
Open Cycle 2
Propulsion SystemGas Core NTR
Exhaust Velocity50,000 m/s
Specific Impulse5,097 s
Thrust5,000,000 N
Thrust Power0.1 TW
Mass Flow100 kg/s
FuelFission:
Uranium Hexafluoride
RemassLiquid Hydrogen
Specific Power2 kg/MW
Engine mass30-200 tonne
T/W 17.0 to 2.5
Open Cycle 3
Thrust Power GW
Exhaust velocity25,000 to 69,000 m/s
Thrust19,600 to 108,000 n
Engine mass40 to 110 tonne
T/W0.05 to 0.10
Operating Pressure400 to 2000 atm
Open Cycle MAX
Exhaust Velocity98,000 m/s
Specific Impulse9,990 s
Thrust3,000,000 N
Thrust Power0.15 TW
Mass Flow31 kg/s
Total Engine Mass15,000 kg
T/W20
FuelFission:
Uranium Hexafluoride
Tanker
Propulsion SystemGas Core NTR
Open-cycle
Exhaust Velocity35,316 m/s
Specific Impulse3,600 s
Thrust3,500,000 N
Thrust Power61.8 GW
Mass Flow99 kg/s
FuelFission:
Uranium Hexafluoride
RemassLiquid Hydrogen
Wet Mass433,000 kg
Dry Mass268,000 kg
Mass Ratio1.62 m/s
ΔV16,943 m/s

Gaseous core fission / nuclear thermal rocket AKA consumable nuclear rocket, plasma core, fizzer, cavity reactor rocket. The limit on NTR-SOLID exhaust velocities is the melting point of the reactor. Some engineer who obviously likes thinking "outside of the box" tried to make a liability into a virtue. They asked the question "what if the reactor was already molten?"

Gaseous uranium is injected into the reaction chamber until there is enough to start a furious chain reaction. Hydrogen is then injected from the chamber walls into the center of this nuclear inferno where is 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.

##### Coaxial
Coaxial
Exhaust Velocity17,658 m/s
Specific Impulse1,800 s
Thrust17,800,000 N
Thrust Power0.2 TW
Mass Flow1,008 kg/s
Total Engine Mass127,000 kg
T/W14
FuelFission:
Uranium Hexafluoride
ReactorGas Core
Open-Cycle
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power1 kg/MW
NASA-Lewis
Thrust Power0.495GW
Exhaust velocity22,000 m/s
Specific Impulse2,200 s
Thrust45,000 n
Engine mass66,000 kg
T/W0.68
Diameter5 m
Length5 m
Fuel Temp20,000° R
Propellant Temp10,000° R

Gaseous core coaxial flow fission / nuclear thermal rocket.

The basic problem of gas core nuclear rockets is ensuring that the hot propellant escapes from the exhaust nozzle, but the nuclear fuel does not. In this concept, the propellant and fuel are kept separate by a velocity differential. That is, a central, slow moving stream of fission fuel heats an annular, fast moving stream of hydrogen.

Yes, the uranium jet is aimed straight at the exhaust nozzle. But they figured the uranium loss would be acceptable as long as 25 to 50 times as much hydrogen propellant escapes compared to uranium fuel (measured by mass).

No, the concept does not work very well. In theory the difference in velocity should keep the uranium/plutonium and the hydrogen separate. Unfortunately the velocity differential at the boundry between the propellant and fuel generates shear forces. The fast hydrogen strips off uranium atoms from the slow fuel plume like a carpenter's plane (laminar and turbulent mixing processes). This means the hydrogen to uranium escape ratio drops below 25.

The concept seems to have been abandoned.

##### Wheel Flow

This is from Wheel-flow gaseous-core reactor concept (1965).

John Evvard figures this gas-core rocket will have (like all the others) an upper limit of about 3,000 seconds of specific impulse, exhaust velocity of about 29,400 m/s. The design is trying to increase the propellant to fuel mass flow ratio to something between 25 and 50. Since uranium has something like 238 times the molecular weight of hydrogen increasing the mass flow ratio is very hard to do.

The brute-force approach does not work. If you increase the engine pressure to 2,000 psi with a partial-pressure ratio of 80, preventing the reaction chamber from exploding will increase the reactor mass to something between 250,000 to 500,000 pounds. With that penalty weight the propellant load will have to exceed 500,000 to 1,000,000 pounds to capitalize on the increase specific impulse the engine enjoys over a conventional solid-core NTR. And even then the fuel mass flow ratio would be below 25. So this is a dead end.

So the standard solution is to somehow make an incredibly high hydrogen-uranium volume flow ratio.

There are numerous schemes to increase the volume flow.

The vortex-confined GCR makes a vortex of gaseous uranium (sort of a smoke ring) with the center hole aligned with the thrust axis. Hydrogen is injected around the outer edge of the vortex, travels radially across the furiously fissioning uranium being heated all the way, enters the hole in the center of the smoke ring, turns 90 degrees and goes rushing out of the hole and out of the exhaust nozzle.

The pious hope was that the centrifugal forces acting on the heavier uranium atoms would counteract the diffusion drag of the inwardly moving hydrogen. Sadly the drag produced by the flowing hydrogen is so great that it carries along too much of the valuable uranium.

The coaxial-flow reactor was another idea that failed even harder. The uranium gas in the center moved really slow while the hydrogen gas around the rim moved really fast. The regrettable result was the velocity difference caused shear forces which allowed the dastardly hydrogen to drag uranium along with it right out the exhaust nozzle.

John Evvard had a fresh idea: the Wheel-Flow Confined GCR.

The problem with the vortex confined GCR was that the hydrogen moves through the uranium. This allows the hydrogen to drag along some uranium. The problem with the coaxial-flow is that though the hydrogen doesn't move through the uranium, it is moving at a vastly different velocity. This causes shear forces that allow the hydrogen to drag along some uranium.

So Evvard tried to find a geometry where the hydrogen does not move through the uranium and it moves at the same velocity as the uranium.

In the Wheel-flow there is a cylinder of gaseous fissioning uranium in the center of the chamber, spinning around its long axis.

Hydrogen is injected at the outer surface of the cylinder and moves along the surface, not moving through the uranium. This avoid the vortex-confined GCR's problem. The hydrogen moves at the same velocity as the uranium gas cylinder. This avoids the coaxial-flow GCR's problem. The uranium and the hydrogen rotate as one, as if they were a solid wheel.

After one rotation of the cylinder the hydrogen is good and hot. It then exits tangentally from the chamber into an array of exhaust nozzles. And there is your thrust.

Uranium will be lost due to fission and some unavoidable diffusion into the hydrogen. Fresh uranium will be injected from the two end walls, entering the long axis of the uranium cylinder. The end wall will also rotate to match the wheel, to avoid stirring up turbulence.

The main drawback is that the boundary layer between the hydrogen and uranium is unstable. Any blob of uranium entering the hydrogen blanket will be accelerated outward by simple boyancy. This could possibly be stablized by an axial magnetic field. The fissioning uranium is more ionized than the hydrogen so the magentic field will grab the uranium more firmly.

Since the temperature inside the reaction chamber is hot enough to vaporize any material object the ions are moving like microscopic bats from hell. You'd think the high uranium molecular velocities would make the uranium cloud instantly explode to fill the chamber. Luckily the mean free path of individual atoms is a microscopic 10-7 meters or less (one micrometer, about the length of a bacteria). Since the hot uranium atoms cannot move further than the span of a typical e coli germ without crashing into other atoms their effective speed is slowed down about the same as the wheel rotational velocity.

The report is a little vague about this design. It says that if the wheel-flow engine is used in a gravitational field, the spinning cylinder of fissioning uranium might settle to the bottom of the chamber, which is bad. However, unless you were using an open-cycle gas core nuclear engine spraying radioactive death from the nozzle as an aircraft engine I don't see the application.

The idea seems to be that while some hydrogen is injected around the uranium gas cylinder for coolant, most of the propellant hydrogen goes across the top along the axial line. I guess the propellant lowers the gas pressure enough to levitate the uranium cylinder.

In the standard wheel design, the end walls will have to be cooled since they are exposed to the fury of fissioning uranium. This can be avoided by bending the uranium gas cylinder so the ends meet, converting the cylinder into a torus donut shape. Since it is now a ring there are no end walls and no need to cool them.

The problem is that the end walls were where the fresh uranium was injected, and it is unclear how to refresh the torus.

This design makes a bit more sense. It uses a torus of uranium gas. The rocket rotates around the thrust axis to make artificial gravity. This pulls the torus outward, making it expand. Meanwhile the propellant hydrogen is roaring down the thrust axis, being heated and expelled out the exhaust nozzle. This lowers axial gas pressure and pulls the torus inward, making it contract.

Between the artificial gravity and Bernoulli's principle the torus of uranium is held in place.

Of course there is still the unsolved problem of how to refresh the torus.

##### Nuclear Salt Water
NSWR
20% UTB
Exhaust Velocity66,000 m/s
Specific Impulse6,728 s
Thrust12,900,000 N
Thrust Power425.7 GW
Mass Flow195 kg/s
Total Engine Mass33,000 kg
T/W40
FuelFission:
Uranium Tetrabromide
ReactorGas Core
Open-Cycle
RemassWater
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorPusher Plate
Specific Power0.08 kg/MW
90% UTB
Exhaust Velocity4,700,000 m/s
Specific Impulse479,103 s
Thrust13,000,000 N
Thrust Power30.6 TW
Mass Flow3 kg/s

This concept by Dr. Zubrin is considered far-fetched by many scientists. The fuel is a 2% solution of 20% enriched Uranium Tetrabromide in water. A Plutonium salt can also be used.

Just to make things clear, there are two percentages here. The fuel is a 2% solution of uranium tetrabromide and water. That is, 2 molecules of uranium tetrabromide per 100 molecules of water.

But the uranium tetrabromide can be 20% enriched. This means that out of every 100 atoms of uranium (or molecules of uranium tetrabromide), 20 are fissionable Uranium-235 and 80 are non-fissionable uranium. If it is 90% enriched, then 90 atoms are Uranium-235 and 10 atoms are non-fissionable. As a side note, 90% enriched is considered "weapons-grade".

The fuel tanks are a bundle of pipes coated with a layer of boron carbide neutron damper. The damper prevents a chain reaction. The fuel is injected into a long cylindrical plenum pipe of large diameter, which terminates in a rocket nozzle. Free of the neutron damper, a critical mass of uranium soon develops. The energy release vaporizes the water, and the blast of steam carries the still reacting uranium out the nozzle.

It is basically a continuously detonating Orion type drive with water as propellant. Although Zubrin puts it like this:

He also notes that it is preferable to subject your spacecraft to a steady acceleration (as with the NSWR) as opposed to a series of hammer-blow accelerations (as with Orion).

The controversy is over how to contain such a nuclear explosion. Zubrin maintains that skillful injection of the fuel can force the reaction to occur outside the reaction chamber. He says that the neutron flux is concentrated on the downstream end due to neutron convection. Other scientists are skeptical.

Naturally in such a spacecraft, damage to the fuel tanks can have unfortunate results (say, damage caused by hostile weapons fire). Breach the fuel tubes and you'll have a runaway nuclear chain reaction on your hands. Inside your ship.

The advantage of NSWR is that this is the only known propulsion system that combines high exhaust velocity with high thrust (in other words, it is a Torchship). The disadvantage is that it combines many of the worst problems of the Orion and Gas Core systems. For starters, using it for take-offs will leave a large crater that will glow blue for several hundred million years, as will everything downwind in the fallout area.

Zubrin calculates that the 20% enriched uranium tetrabromide will produce a specific impulse of about 7000 seconds (69,000 m/s exhaust velocity), which is comparable to an ion drive. However, the NSWR is not thrust limited like the ion drive. Since the NSWR vents most of the waste heat out the exhaust nozzle, it can theoretically produce jet power ratings in the thousands of megawatts (meaning it is not power limited, like other nuclear propulsion). Also unlike the ion drive, the engine is relatively lightweight, with no massive power plant required.

Zubrin suggests that a layer of pure water be injected into the plenum to form a moving neutron reflector and to protect the plenum walls and exhaust nozzle from the heat. One wonders how much protection this will offer.

Zubrin gives a sample NSWR configuration. It uses as fuel/propellant a 2% (by number) uranium bromide aqueous solution. The uranium is enriched to 20% U235. This implies that B2 = 0.6136 cm-2 (the material buckling, equal to vΣfa)/D) and D = 0.2433 cm (diffusion coefficent).

Radius of the reaction plenum is set to 3.075 centimeters. this implies that A2 = 0.6117 cm-2 and L2 = 0.0019. Since exponential detonation is desired, k2 = 2L2 = 0.0038 cm-2. Then k = U / 2D = 0.026 cm-1 and U = 0.03.

If the velocity of a thermal neutron is 2200 m/s, this implies that the fluid velocity needs to be 66 m/s. This is only about 4.7% the sound speed of room temperature water so it should be easy to spray the fuel into the plenum chamber at this velocity.

The total rate of mass flow through the plenum chamber is about 196 kg/s.

Complete fission of the U235 would yield about 3.4 x 1012 J/kg. Zubrin assumes a yield of 0.1% (0.2% at the center of the propellant column down to zero at the edge), which would not affect the material buckling during the burn. This gives an energy content of 3.4 x 109 J/kg.

Assume a nozzle efficiency of 0.8, and the result is an exhaust velocity of 66,000 m/s or a specific impules of 6,7300 seconds. The total jet power is 427 gigawatts. The thrust is 12.9 meganewtons. The thrust-to-weight ratio will be about 40, which implies an engine mass of about 33 metric tons.

For exponetial detonation, kz has to be about 4 at the plenum exit. Since k = 0.062 cm-1, the plenum will have to be 65 cm long. The plenum will be 65 cm long with a 3.075 cm radius, plus an exhaust nozzle.

Zubrin then goes on to speculate about a more advanced version of the NSWR, suitable for insterstellar travel. Say that the 2% uranium bromide solution used uranium enriched to 90% U235 instead of only 20%. Assume that the fission yield was 90% instead of 0.1%. And assume a nozzle efficency of 0.9 instead of 0.8.

That would result in an exhaust velocity of a whopping 4,725,000 m/s (about 1.575% c, a specific impulse of 482,140 seconds). In a ship with a mass ratio of 10, it would have a delta V of 3.63% c. Now you're talkin...

### Fission Fragment

#### Fission Fragment

George Chapline
Exhaust velocity980,000 m/s

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 5% the speed of light! 15,000,000 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 through a reactor core, where atoms of nuclear fuel would undergo fission. The fission fragments would be directed by magnetic fields into an exhaust beam.

The drawback of this design is that too many of the fragments fail to escape the fuel coat (which adds no thrust but does heat up the coat) and too many hit the carbon filaments (which adds no thrust but does heat up the filaments). This is why the disks spin at high speed, otherwise they'd melt.

Dusty Plasma
550AU
Thrust22 N
Thrust Power0.2 GW
Mass Flow1.00e-06 kg/s
T/W2.49e-04
Specific Power55 kg/MW
0.5LY
Thrust344 N
Thrust Power2.6 GW
Mass Flow2.30e-05 kg/s
T/W4.00e-03
Specific Power3 kg/MW
All
Exhaust Velocity15,000,000 m/s
Specific Impulse1,529,052 s
Total Engine Mass9,000 kg
FuelFission:
Uranium 235
ReactorGas Core
MHD Choke
RemassReaction
Products
Remass AccelFission-Fragment
Thrust DirectorMagnetic Nozzle

Rodney Clark and Robert Sheldon solve the problem with their Dusty plasma bed reactor (report).

You take the fission fuel and grind it into dust grains with an average size of 100 nanometers (that is, about 1/20th the thickness of the fuel coating in dr. Chapline's design). This does two things [A] most of the fragments escape and [B] the dust particles have such a high surface to volume ratio that heat (caused by fragments which fail to escape) readily dissipates, preventing the dust particles from melting.

The dust is suspended in the center of a reaction chamber whose walls are composed of a nuclear moderator. Power reactors will use beryllium oxide (BeO) as a moderator, but that is a bit massive for a spacecraft. The ship will probably use lithium hydride (LiH) for a moderator instead, since is only has one-quarter the mass. Probably about six metric tons worth. The dust is suspended electrostatically or magnetically by a containment field generator. The dust is heated up by radio frequency (RF) induction coils. The containment field generator will require superconductors, which will probably require a coolant system of its own.

The dust particles are slow and are relatively massive, while the fission fragments are fast and not very massive at all. So the magnetic field can be tailored so it holds the dust but allows the fission fragments to escape. Magnetic mirrors ensure that fragments headed the wrong way are re-directed to the exhaust port.

One valuable trick is that you can use the same unit for thrust or to generate electricity. Configure the magnetic field so that the fragments escape "downward" through the exhaust port and you have thrust. Flip a switch to change the magnetic field so that the fragments escape upward into deceleration and ion collection electrodes and you generate electricity. As a matter of fact, it is so efficient at generating electricity that researchers are busy trying to adapt this for ground based power plants. But I digress.

The dust is only sufficient for a short period of critical nuclear reaction so it must be continuously replenished. The thermal energy released by fission events plus heat from collisions between fission fragments and dust grains create intense heat within the dust cloud. Since there is no core cooling flow, the reactor power is limited to the temperature at which the dust can radiatively cool itself without vaporizing. The interior of the reaction chamber walls will protected by a mirrored (95% reflection) heat shield attached to a heat radiator. The outer moderator layer will have its own heat shield.

Clark and Sheldon roughed out a propulsion system. It had six tons for the moderator, 2 tons for radiators and liquid metal cooling, 1 ton for magnets, power recovery, and coils, for a grand total of 9 tons. The reaction chamber will be about 1 meter in diameter and 10 meters long. The moderator blanket around the chamber will be about 40 centimeters thick. The thrust is a function the size of the cloud of fissioning dust, and is directly related to the power level of the reactor. There is a limit to the maximum allowed power level, set by the coolant system of the reaction chamber. Clark and Sheldon estimate that only about 46% of the fission fragments provide thrust while the rest are wasted. See the report for details.

In the table, the 550AU engine is for a ten year journey to the Solar gravitational lensing point at 550 astronomical units (so you can use the sun as a giant telescope lens). The 0.5LY engine is for a thirty year trip to the Oort cloud of comets. These are constant acceleration brachistochrone trajectories, the 550AU mission will need a reactor power level of 350 MW and the 0.5LY mission will need 5.6 GW. Don't forget that the engine power is only 46% efficient, that's why the table thrust values are lower.

Werka FFRE
First Generation
Exhaust Velocity5,170,000 m/s
Specific Impulse527,013 s
Thrust43 N
Thrust Power0.1 GW
Mass Flow8.00e-06 kg/s
Total Engine Mass113,400 kg
T/W3.90e-05
FuelFission:
Plutonium 239
ReactorGas Core
MHD Choke
RemassReaction
Products
Remass AccelFission-Fragment
Thrust DirectorMagnetic Nozzle
Specific Power1,020 kg/MW
HOPE FFRE
Propulsion SystemWerka FFRE
Wet Mass303,000 kg
Dry Mass295,000 kg
Mass Ratio1.03 m/s
ΔV138,336 m/s

Robert Werka has a more modest and realistic design for his fission fragment rocket engine (FFRE). He figures that a practical design will have an exhaust velocity of about 5,200,000 m/s instead of his estimated theoretical maximum of 15,000,000 m/s. His lower estimate is still around 1.7% the speed of light so we are still talking about sub 2.0 mass ratios. Collisions between fission fragments and the dust particles is responsible for the reduction in exhaust velocity.

Incidentally the near relativistic exhaust velocity reduces radioactive contamination of the solar system. The particles are traveling well above the solar escape velocity (actually they are even faster than the galactic escape velocity) so all the radioactive exhaust goes shooting out of the solar system at 0.017c.

The dusty fuel is nanometer sized particles of slightly critical plutonium carbide, suspended and contained in an electric field. A moderator of deuterated polyethylene reflects enough neutrons to keep the plutonium critical, while control rods adjust the reaction levels. The moderator is protected from reaction chamber heat by a heat shield, an inner layer composed of carbon-carbon to reflect infrared radiation back into the core. The heat shield coolant passes through a Brayton cycle power generator to create some electricty, then the coolant is sent to the heat radiator.

The details of Werka's initial generation FFRE can be found in the diagram below. The reaction chamber is about 5.4 meters in diameter by 2.8 meters long. The magnetic nozzle brings the length to 11.5 meters. The fuel is uranium dioxide dust which melts at 3000 K, allowing a reactor power of 1.0 GW. It consume about 29 grams of uranium dioxide dust per hour (not per second). Of the 1.0 GW of reactor power, about 0.7 GW of that is dumped as waste heat through the very large radiators required.

The second most massive component is the magnetic mirror at the "top" of the reaction chamber. Its purpose is to reflect the fission fragments going the wrong way so they turn around and travel out the exhaust nozzle. Surrounding the "sides" of the reaction chamber is the collimating magnet which directs any remaining wrong-way fragments towards the exhaust nozzle. The exhaust beam would cause near-instantaneous erosion of any material object (since it is electrically charged, relativistic, radioactive grit). It is kept in bounds and electrically neutralized by the magnetic nozzle cage.

#### Afterburner Fission Fragment

AFFRE
EngineAFFRE
Engine Mass
(reactor)
107,000 kg
Engine Mass
(mod oil)
91,000 kg
Engine Mass
(total)
268,961 kg
Reactor Power2.5 GW
Thrust4,651 N
Thrust Power730 MW
Specific
Impulse
32,000 sec
Exhaust
Velocity
313,900 m/s
Mass Flow
(FF)
3.12×10-5 kg/s
Mass Flow
(Hydrogen)
0.0179 kg/s
Mass Flow
(Total)
0.018 kg/s
T/W0.002

Robert Werka has apparently figured out a new configuration for his fission-fragment rocket engine (FFRE).

As with most engines that have high specific impulse and exhaust velocity, the thrust of a FFRE is pathetically small. Ah, but there is a standard way of dealing with this problem: shifting gears. What you do is inject cold propellant into the exhaust ("afterburner"). The fission fragment exhaust loses energy while the cold propellant gains energy. The combined exhaust velocity of the fission fragment + propellant energy is lower than the original pure fission fragment, so the specific impulse goes down. However the propellant mass flow goes up since the combined exhaust has more mass than the original pure fission fragment. So the thrust goes up.

Now you have an Afterburner fission-fragment rocket engine (AFFRE).

As you are probably tired of hearing, this means the engine has shifted gears by trading specific impulse for thrust.

Shifting Gears
EngineIspThrust
FFRE527,000 sec43 Newtons
AFFRE32,000 sec4,651 Newtons

The heart of the engine is a standard "dusty plasma" fission fragment engine. A cloud of nanoparticle-sized fission fuel is held in an electrostatic field inside a neutron moderator. Atoms in the particles are fissioning like crazy, spewing high velocity fission products in all directions. These become the exhaust, directed by a magnetic nozzle.

The AFFRE alters this a bit. Instead of a cylindrical reactor core it uses half a torus. Each end of the torus has its own magnetic nozzle. But the biggest difference is that cold hydrogen propellant is injected into the flow of fission fragments as an afterburner, in order to shift gears.

In the diagram above, the magnetic nozzles are the two frameworks perched on top of the reactor core. It is a converging-diverging (C-D) magnetic nozzle composed of a series of four beryllium magnetic rings (colored gold in the diagram). Note how each frame holding the beryllium rings is shaped like an elongated hour-glass, that is the converting-diverging part. The fission fragment plume emerges from the reactor core, is squeezed (converges) down until it reaches the midpoint of the magnetic nozzle, then expands (diverges) as it approaches the end of the nozzle. At the midpoint is the afterburner, where the cold hydrogen propellant is injected.

The semi-torus has a major and minor radius of 3 meters. The overall length of the engine is 13 meters. The reactor uses 91 metric tons of hydrocarbon oil as a moderator. This means the heavy lift vehicle can launch the engine "dry" with no oil moderator. In orbit the oil moderator can be easily injected into the reactor, at least easier than building the blasted thing in free fall out of graphite bricks.

Fission Sail

#### Antimatter-Driven Sail

The sail is made of graphite and carbon-carbon fiber, infused with a tiny amount of uranium. It is subjected to a misting of antiprotons. These induce uranium atoms to fission, with the recoil pushing the sail. Since this is nuclear powered, the sail does not have to be kilometers in diameter, five meters will do. 30 miligrams of antiprotons could push the sail to the Kuiper Belt.

## Pulse

### Orion

Fission Orion
Exhaust Velocity43,000 m/s
Specific Impulse4,383 s
Thrust263,000 N
Thrust Power5.7 GW
Mass Flow6 kg/s
Total Engine Mass200,000 kg
T/W0.13
FuelFission:
Uranium 235
ReactorPulse Unit
RemassTungsten
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorPusher Plate
Specific Power35 kg/MW
Fusion Orion
Exhaust Velocity73,000 m/s
Specific Impulse7,441 s
Thrust292,000 N
Thrust Power10.7 GW
Mass Flow4 kg/s
Total Engine Mass200,000 kg
T/W0.15
FuelD-D Fusion
Specific Power19 kg/MW
1959 Orion 1st Gen
Thrust Power1,600 GW
Exhaust velocity39,000 m/s
Thrust80,000,000 n
Engine mass1,700 tonne
T/W >1.0yes
1959 Orion 2nd Gen
Thrust Power24,000 GW
Exhaust velocity120,000 m/s
Thrust400,000,000 n
Engine mass3,250 tonne
T/W >1.0yes
ORION USAF 10m
Exhaust Velocity32,900 m/s
Specific Impulse3,354 s
Thrust2,000,000 N
Thrust Power32.9 GW
Mass Flow61 kg/s
Total Engine Mass107,900 kg
T/W2
Wet Mass475,235 kg
Dry Mass180,975 kg
Mass Ratio2.63 m/s
ΔV31,763 m/s
Specific Power3 kg/MW
ORION 4K ton battleship
Exhaust Velocity39,000 m/s
Specific Impulse3,976 s
Thrust80,000,000 N
Thrust Power1.6 TW
Mass Flow2,051 kg/s
Total Engine Mass1,700,000 kg
T/W4.80
Specific Power1.09 kg/MW
ΔV 10 km/s
Wet Mass4,000,000 kg
Dry Mass3,100,000 kg
Mass Ratio1.29 m/s
ΔV9,941 m/s
ΔV 21 km/s
Wet Mass4,000,000 kg
Dry Mass2,353,000 kg
Mass Ratio1.70 m/s
ΔV20,694 m/s
ΔV 30 km/s
Wet Mass4,000,000 kg
Dry Mass1,852,000 kg
Mass Ratio2.16 m/s
ΔV30,031 m/s
Exhaust Velocity120,000 m/s
Specific Impulse12,232 s
Thrust400,000,000 N
Thrust Power24.0 TW
Mass Flow3,333 kg/s
Total Engine Mass3,250,000 kg
T/W13
Specific Power0.14 kg/MW
ΔV 10 km/s
Wet Mass10,000,000 kg
Dry Mass9,199,000 kg
Mass Ratio1.09 m/s
ΔV10,019 m/s
ΔV 15.5 km/s
Wet Mass10,000,000 kg
Dry Mass8,772,000 kg
Mass Ratio1.14 m/s
ΔV15,722 m/s
ΔV 20 km/s
Wet Mass10,000,000 kg
Dry Mass8,403,000 kg
Mass Ratio1.19 m/s
ΔV20,880 m/s
ΔV 30 km/s
Wet Mass10,000,000 kg
Dry Mass7,813,000 kg
Mass Ratio1.28 m/s
ΔV29,616 m/s
ΔV 100 km/s
Wet Mass10,000,000 kg
Dry Mass4,348,000 kg
Mass Ratio2.30 m/s
ΔV99,944 m/s
Orion MAX
Exhaust Velocity9,800,000 m/s
Specific Impulse998,981 s
Thrust8,000,000 N
Thrust Power39.2 TW
Mass Flow0.82 kg/s
Total Engine Mass8,000 kg
T/W102
FuelProton-Proton
Fusion
RemassTungsten
Specific Power2.04e-04 kg/MW

Orion AKA "old Boom-boom" is the ultimate consumable nuclear thermal rocket, based on the "firecracker under a tin can" principle. Except the tin can is a spacecraft and the firecracker is a nuclear warhead.

This concept has the spacecraft mounted with shock absorbers on an armored "pusher plate". A stream of small (5 to 15 kiloton) fission or fusion explosives are detonated under the plate to provide thrust. While you might find it difficult to believe that the spacecraft can survive this, you will admit that this will give lots of thrust to the spacecraft (or its fragments). On the plus side, a pusher plate that can protect the spacecraft from the near detonation of nuclear explosives will also provide dandy protection from any incoming weapons fire. On the minus side I can hear the environmentalists howling already. It will quite thoroughly devastate the lift-off site, and give all the crew bad backs and fallen arches. And they had better have extra-strength brassieres and athletic supporters.

Mathematician Richard Courant viewed an Orion test and said "Zis is not nuts, zis is super-nuts."

This section is about the Orion propulsion system. If you want all the hot and juicy details about various versions of Orion spacecraft go here.

Please note that Orion drive is pretty close to being a torchship, and is not subject to the Every gram counts rule. It is probably the only torchship we have the technology to actually build today.

If you want the real inside details of the original Orion design, run, do not walk, and get a copies the following issues of of Aerospace Projects Review: Volume 1, Number 4, Volume 1, Number 5, and Volume 2, Number 2. They have blueprints, tables, and lots of never before seen details.

If you want your data raw, piled high and dry, here is a copy of report GA-5009 vol III "Nuclear Pulse Space Vehicle Study - Conceptual Vehicle Design" by General Atomics (1964). Lots of charts, lots of graphs, some very useful diagrams, almost worth skimming through it just to admire the diagrams.

The sad little secret about Orion is that the mission it is best suited for is boosting heavy payloads into orbit. Which is exactly the mission that the enviromentalist and the nuclear test ban treaty will prevent. Orion has excellent thrust, which is what you need for lift-off and landing. Unfortunately its exhaust velocity is pretty average, which is what you need for efficient orbit-to-orbit maneuvers.

Having said that, there is another situation where high thrust is desirable: a warship jinking to make itself harder to be hit by enemy weapons fire. It is also interesting to note that the Orion propulsion system works very well with the bomb-pumped laser weapons system.

Each pulse unit is a tiny nuclear bomb, encased in a "radiation case" that has a hole in the top. A nuclear blasts is initially mostly x-rays. The radiation case is composed of a material that his opaque to x-rays (depleted uranium). The top hole thus "channels" the flood of x-rays in an upwards direction (at least in the few milliseconds before the bomb vaporizes the radiation case).

The channeled x-rays then strike the "channel filler" (beryllium oxide). The channel filler transforms the atomic fury of x-rays into an atomic fury of heat.

Lying on top of the channel filler is the disc of propellant (tungsten). The atomic fury of heat flashes the tungsten into a jet of ionized tungsten plasma, traveling at high velocity (in excess of 1.5 × 105 meters per second). This crashes into the pusher plate, accelerating the spacecraft. It crashes hard. You will note that there are two stages of shock absorbers between the pusher plate and the spacecraft, preventing instant crew death.

The ratio of beryllium oxide to tungsten is 4:1.

The thickness of the beryllium oxide and tungsten should be such to serve as a shield to protect the engine and upper vehicle from the neutron and high-energy gamma radiation produced by the nuclear explosion. This sets a lower limit on the thickness of the propellant and channel filler for a particular design.

The jet is confined to a cone about 22.5 degrees (instead of in all directions). The detonation point is positioned such that the 22.5 cone exactly covers the diameter of the pusher plate. The idea is that the wider the area of the cone, the more spread out the impulse will be, and the larger the chance that the pusher plate will not be utterly destroyed by the impulse.

It is estimated that 85% of the energy of the nuclear explosion can be directed in the desired direction. The pulse units are popped off at a rate of about one per second. A 5 kiloton charge is about 1,152 kg. The pulses are so brief that there is no appreciable "neutron activation", that is, the neutron from the detonations do not transmute parts of the spacecraft's structure into radioactive elements. This means astronauts can exit the spacecraft and do maintenance work shortly after the pulse units stop detonating.

The device is basically a nuclear shaped charge. A pulse unit that was not a shaped charge would of course waste most of the energy of the explosion. Figure that 1% at best of the energy of a non-shaped-charge explosion would actually hit the pusher plate, what a waste of perfectly good plutonium.

Each charge accelerates the spacecraft by roughly 12 m/s. A 4,000 ton spacecraft would use 5 kiloton charges, and a 10,000 ton spacecraft would use 15 kiloton charges. For blast-off, smaller charges of 0.15 kt and 0.35 kt respectively would be used while within the Terra's atmosphere. The air between the charge and the pusher plate amplifies the impulse delivered (it is extra propellant), so if you are not in airless space you can get away with a smaller kt yield.

How much weapons-grade plutonium will each charge require? As with most details about nuclear explosives, specifics are hard to come by. According to GA-5009 vol III , pulse units with 2.0×106 newtons to 4.0×107 newtons all require approximately 2 kilograms per pulse unit, with 1964 technology. It goes on to say that advances in the state of the art could reduce the required amount of plutonium by a factor 2 to 4, especially for lower thrust units. 2.0×106 n is 1 kiloton, I'm not sure what 4.0×107 n corresponds to, from the document I'd estimate it was about 15 kt. Presumably the 2 kg plutonium lower limit is due to problems with making a critical mass, you need a minimum amount to make it explode at all.

According to Scott Lowther, the smallest pulse units were meant to propel a small ten-meter diameter Orion craft for the USAF and NASA. The units had a yield ranging from one-half to one kiloton. The USAF device was one kiloton, diameter 36 centimeters, mass of 86 kilograms, tungsten propellant mass of 34.3 kilograms, jet of tungsten plasma travels at 150,000 meters per second. One unit would deliver to the pusher plate a total impulse of 2,100,000 newton-seconds. Given the mass of the ten-meter Orion, detonating one pulse unit per second would give an acceleration well over one gee. According to my slide rule, this implies that the mass of the ten-meter Orion is a bit under 210 metric tons.

Pulse UnitYieldMassDia.HeightPropellant
(percent)
Det.
Interval
Propellant
Velocity
Effective
Exhaust
Velocity
(Isp)
Thrust
per unit
Effective
Thrust
NASA 10m Orion
(vacuum)
141 kg0.86 s18,200 m/s
(1,850 s)
3.0×106 N3.5×106 N
USAF 10m Orion
(vacuum)
1 kt79 kg
(86 kg?)
0.33 m0.61 m34.3 kg
(40%)
1 s1.5×105 m/s25,800 m/s
(2,630 s)
2.0×106 N2.0×106 N
20m Orion
(vacuum)
450 kg0.87 s30,900 m/s
(3,150 s)
1.4×107 N1.6×107 N
4000T Orion
(atmo)
0.15 kt1,152 kg0.81 m0.86 m1.1 s1.17×105 m/s42,120 m/s
(4,300 s)
8.8×107 N8.0×107 N
4000T Orion
(vacuum)
5 kt1,152 kg0.81 m0.86 m415 kg
(36%)
1.1 s1.17×105 m/s42,120 m/s
(4,300 s)
8.8×107 N8.0×107 N
10,000T Orion
(atmo)
0.35 kt118,000 m/s
(12,000 s)
4.0×108
10,000T Orion
(vacuum)
15 kt118,000 m/s
(12,000 s)
4.0×108
20,000T Orion
(vacuum)
29 kt1,150 kg0.8 m
• Pulse Unit: The type of Orion spacecraft that uses this unit, and whether it is an atmospheric or vacuum type.
• Yield: Nuclear explosive yield (kilotons)
• Mass: Mass of the pulse unit
• Dia.: Diameter of pulse unit
• Height: Height of pulse unit
• Propellant (percent): Mass of tungsten propellant in kilograms, as percentage of pulse unit mass in parenthesis.
• Det. Interval: Time delay interval between pulse unit detonations.
• Propellant Velocity: The velocity the tungsten propellant plasma travels at. Do not use this for delta V calculations.
• Effective Exhaust Velocity (Isp): A value for exhaust velocity suitable for delta V calculations. Specific impulse in parenthesis.
• Thrust per unit: Amount of thrust produced by detonating one pulse unit.
• Effective Thrust: Thrust per second. Calculated by taking Thrust per unit and dividing by Det. Interval.

Tungsten has an atomic number (Z) of 74. When the tungsten plate is vaporized, the resulting plasma jet has a relatively low velocity and diverges at a wide angle (22.5 degrees). Now, if you replace the tungsten with a material with a low Z, the plasma jet will instead have a high velocity at a narrow angle. The jet angle also grows narrower as the thickness of the plate is reduced. This makes it a poor propulsion system, but an effective weapon. Instead of a wall of gas hitting the pusher plate, it is more like a directed energy weapon. The military found this to be fascinating, who needs cannons when you can shoot spears of pure nuclear flame? The process was examined in a Strategic Defense Initiative project called "Casaba-Howitzer", which apparently is still classified. Which is not surprising but frustrating if one is trying to write a science fiction novel or spacecraft combat game.

NASA has been quietly re-examining ORION, under the new name of "External Pulsed Plasma Propulsion". As George Dyson observed, the new name removes most references to "Nuclear", and all references to "Bombs."

For details about spacecraft using Orion propulsion, go here.

Oh, and another thing. ORION is fantastic for boosting unreasonably huge payloads into orbit and it is pretty great for orbit to orbit propulsion. But trying to use it to land is not a very good idea. At least not on a planet with an atmosphere.

#### Orion Thrust and Isp

Even though only a fraction of the pulse unit's mass is officially tungsten propellant, you have to count the entire mass of the pulse unit when figuring the mass ratio. The mass of the Orion spacecraft with a full load of pulse units is the wet mass, and the mass with zero pulse units is the dry mass.

The thrust is not applied constantly, it is in the form of pulses separated by a fixed detonation interval. Generally the interval is from about half a second to 1.5 seconds. This means to figure the "effective" thrust you take the thrust-per-pulse-unit and divide it by the detonation interval in seconds. So if each pulse unit gives 2×106 Newtons, and they are detonated at 0.8 second intervals, the effective thrust is 2×106 / 0.8 = 2.5×106 Newtons

Obviously the converse is if you have the effective thrust, you multiply it by the detonation interval to find the thrust-per-pulse-unit. So if the effective thrust is 3.5×106 N and the units are detonated at 0.86 second intervals, the thrust-per-pulse-unit is 3.5×106 N * 0.86 = 3.01×106 Newtons

There are some interesting equations in GA-5009 vol III on pages 25 and 26 on the subject of nuclear pulse units. These were developed in the study for the 10 and 20 meter NASA Orion spacecraft, and they heavily rely upon a number of simplifying assumptions. These were for first generation pulse units, with the assumption that second generation units would have better performance. So take these with a grain of salt.

These equations are only considered valid over the range 3×106 < FE < 2×108

You are given the amount of thrust you want to get out of the propulsion system: FE and the detonation interval time Dp. From those you calculate the amount of thrust each pulse unit has to deliver Fp:

Fp = FE / Dp

From this the specific impulse, nuclear yield, and the mass of the Orion propulsion module.

Isp = 1 / ((5.30×102 / (Fp * (1 + (2.83×10-3 * Fp1/3)))) + ((4.32×10-2 * (1 + (2.83×10-3 * Fp1/3))) / Fp1/3))

Ve = Isp * g0

Y = 9.30×10-10 * Fp4/3

ME = Fp / (3.6 * g0)

where

• FE = effective thrust (newtons)
• Dp = delay between pulses (seconds)
• Fp = thrust per pulse (newtons)
• Isp = effective specific impulse (seconds)
• Ve = exhaust velocity (m/s)
• Y = size of nuclear yield in pulse unit (kilotons)
• ME = mass of Orion propulsion module (kg)
• g0 = acceleration due to gravity = 9.81 m/s2
• x1/3 = cube root of x

The results are close but do not exactly match the values given in the document, but they are better than nothing

NASA 10-meter Orion
Given Effective Thrust3.5×106 N
Given Detonation Delay0.86 s
ParameterDocument
Value
Equation
Value
Specific Impulse1,850 s1,830 s
Yield1 kt0.4 kt
Propulsion module mass90,946 kg85,245 kg
NASA 20-meter Orion
Given Effective Thrust1.6×107 N
Given Detonation Delay0.87 s
ParameterDocument
Value
Calculated
Value
Specific Impulse3,150 s3,082 s
Yield5 kt3.1 kt
Propulsion module mass358,000 kg394,223 kg

For more in depth calculations of an Orion rocket's specific impulse, read page 1 and page 2. But be prepared for some heavy math.

#### Orion Environmental Impact

Naturally, some people freak out when you tell them about a rocket that rises into orbit by detonating Two! Hundred! Atom! Bombs!. But it actually isn't quite as bad as it sounds.

First off, these are teeny-tiny atom bombs, honest. The nuclear pulse units used in space will be about one kiloton each, while the Nagasaki device was more like 20 kt. And in any event, the nuclear pulse units used in the atmosphere are only 0.15 kt ( about 1/130th the size of the Nagasaki device). This is because the atmosphere converts the explosion x-rays into "blast", increasing the effectiveness of the pulse unit so you can lower the kilotonnage.

So we are not talking about zillions of 25 megaton city-killer nukes scorching the planet and causing nuclear winter.

Some environmentalists howl that Orion should never be used for surface-to-orbit boosts, due to the danger of DUNT-dunt-Dunnnnnnnn Deadly Radioactive Fallout. However, there is a recent report that suggests ways of minimizing the fallout from an ORION doing a ground lift-off (or a, wait for it, "blast-off" {rimshot}). Apparently if the launch pad is a large piece of armor plate with a coating of graphite there is little or no fallout.

By which they mean, little or no ground dirt irradiated by neutrons and transformed into deadly fallout and spread the the four winds.

There is another problem, though, ironically because the pulse units use small low-yield nuclear devices.

Large devices can be made very efficient, pretty much 100% of the uranium or plutonium is consumed in the nuclear reaction. It is much more difficult with low-yield devices, especially sub-kiloton devices. Some of the plutonium is not consumed, it is merely vaporized and sprayed into the atmosphere. Fallout, in other words. You will need to develop low-yield devices with 100% plutonium burn-up, or use fusion devices (with 100% burn-up fission triggers or with laser inertial confinement fusion triggers).

The alternative is boosting the Orion about 90 kilometers up using a non-fallout chemical rocket. Which more or less defeats the purpose of using an Orion engine in the first place. Remember that Orions are best at boosting massive payloads into orbit.

Most of the fallout will fall within 80 kilometers of the launch site. You can also reduce the fallout by a factor of 10 if you launch from near the Magnetic Pole.

When fissionables like plutonium undergo fission, their atoms are split which produces atomic energy. The split atoms are called fission fragments.

The good news is that they have very short half-lives, e.g., in 50 days pretty much all of the Strontium 94 has decayed away (because 50 days is 58,000 St94 half-lives).

The bad news is that they have very short half-lives, this means they are hideously radioactive. Radioactive elements decay by emitting radiation, shorter half-life means more decays per second means a higher dose of radiation per second.

The fragments that come screaming out of the detonation aimed at the sky are no problem. They are moving several times faster than Terra's escape velocity, you will never see them again (Terra's escape velocity is 11.2 km/s, the fragments are travelling like a bat out of hell at 2,000 km/s). The ones aimed towards Terra are a problem. The fragments can be reduced by using fusion instead of fission pulse units. The fragments can also be reduced by designing the pulse units to trade thrust in favor of directing more of the fragments skyward.

A more sophisticated objection to using Orion inside an atmosphere is the sci-fi horror of EMP melting all our computers, making our smart phones explode, and otherwise ruining anything using electricity. But that actually is not much of a problem. EMP is not a concern unless the detonation is larger than one megaton or so, Orion propulsion charges are only a few kilotons (one one-thousandth of a megaton). Ben Pearson did an analysis and concluded that Orion charges would only have EMP effects within a radius of 276 kilometers (the International Space Station has an orbital height of about 370 kilometers). So just be sure your launch site is in a remote location, which you probably would have done anyway.

Naturally watching an Orion blast-off is very bad for your eyes, defined as instant permanent blindness. This is called "eyeburn". While the Orion is below 30 km you definitely need protective goggles or you might be blinded. Above 90 km your eyesight it safe. In between 30 and 90 is the gray area, where prudent people keep their protective goggles on.

Detonating pulse units in space near Terra can create nasty artificial radiation belts. The explosion can pump electrons into the magnetosphere, creating the belt.

There are two factors: detonation altitude from Terra's surface, and magnetic latitude in Terra's magnetic field. If the detonation is within 6,700 kilometers of Terra's surface (i.e., closer than 2 Terran radii from Terra's center) and at a magnetic latitude from 0° to 40°, the radiation belt can last for years. Above 2 Terran radii the radiation belt will last for only weeks, and from latitude 80° to 90°, the radiation belt will last for only a few minutes.

The military discovered this the hard way with the Starfish Prime nuclear test. The instant auroras were very pretty. The instant EMP was very scary, larger than expected (but the test was using a 1.4 megaton nuke, not a 0.001 megaton pulse unit). The artificial radiation belt that showed up a few days later was a very rude surprise. About one-third of all low orbiting satellites were eventually destroyed by the radiation belt.

The radiation belts are harmless to people on Terra, but astronauts in orbit and satellites are at risk.

There are three classes of pulse unit failure modes. Note that in this analysis the USAF had given up and had decided to boost the Orion on top of a chemical rocket.

Typically occurs when the chemical booster burns or explodes on the pad. There will be no nuclear explosion. The pulse units contain chemical explosives, but there is much more explosive potential in the chemical booster fuel. Even if all the pulse units exploded simultaneously there would only be a 1 psi overpressure out to 300 meters and shrapnel hazard out to 2,000 meters.

A chemical booster burn could aerosolize radioactive plutonium from booster units and create a downrange fallout hazard. The solution is to put the launch pad over a pool of water about 10 meters deep. In event of fire, collapse the pad into the pool. The fire would be extinguished and any escaped plutonium will be contained in the water. Many of the pulse units can be recovered and reused.
Class II - Failure to Orbit
The trouble is that the thousands of nuclear pulse units will fall down, probably into uncontrolled territory. As with Class I there will be no nuclear explosion, the chemical explosion will be impressive but not too huge, and there is a danger of radioactive fallout. All in what could very well be a foreign country.

In addition, it will be scattering thousands of containers of weapons grade plutonium in convenient form to cause nuclear weapon proliferation. Or the pulse units could be used as is as impromptu terrorist devices. Though I'm sure the devices will contain fail-safes seven ways to Sunday, the same way nuclear warheads are in order to deal with the possibility of them falling into the Wrong Hands.

Probably the best solution is to command all of the nuclear charges to detonate simultaneously while the spacecraft is at high altitude. This will make one heck of a fireworks display, and may cause an EMP, but nuclear devices in questionable hands is to be avoided at all costs.
Class III - Misfire
If a given pulse unit fails to detonate, the command can be resent repeatably, and/or there can be an automatic on-board destruct system. Otherwise the unit could survive reentry (due to the tungsten propellant plate) causing some damage to the country it hit and causing a foreign policy nightmare to the nation owning the Orion spacecraft.

By about 1963 General Atomic had given up on designing an Orion to lift off from Terra's surface under nuclear power. They put together three plans for using chemical rocket boosters to get the Orion into orbit. Again this is throwing away the big advantage of the Orion, its ability to boost massive payloads.

Mode I
A fully loaded and fully fueled Orion is boosted to an altitude of 90 kilometers and 900 m/s by a chemical rocket. There it stages, and the Orion proceeds into orbit or into mission trajectory under nuclear power. The disadvantage is it requires a subobital start-up of the Orion engine. The Orion engine will need a thrust greater than the mass of the spacecraft, the standard was T/W of 1.25. But high thrust is never a problem with Orion.
Mode II
An empty Orion is loaded with just enough pulse units. It is boosted to an altitude of 90 kilometers and 900 m/s by a chemical rocket. There it stages, and the Orion proceeds into orbit. A second chemical booster rendezvous with the Orion to deliver the payload and a full load of pulse units.
This was the worst plan. It combines the disadvantage of Mode I (by requiring suborbital start-up of the Orion engine) with the disadvantage of Mode III (by requiring orbital assembly).
Mode III
The Orion is boosted into orbit piecemeal as payload on a series of chemical boosters. The Orion is assembled in orbit, then departs on its mission under nuclear power. The main advantage is it avoids the possibility of the entire Orion spacecraft crashing to Terra in the event of a propulsion failure. The second advantage is it allowed a lower thrust Orion unit to be used, but with Orion thrust is never a problem. The main disadvantage is that orbital assembly is time consuming and difficult.

### Zeta-Pinch

Zeta pinch is a type of plasma confinement system that uses an electrical current in the plasma to generate a magnetic field that compresses it. The compression is due to the Lorentz force.

#### Zeta-Pinch Fission

##### Mini-Mag Orion
Mini-Mag Orion
Exhaust Velocity157,000 m/s
Specific Impulse16,004 s
Thrust1,870,000 N
Thrust Power0.1 TW
Mass Flow12 kg/s
Total Engine Mass199,600 kg
T/W0.95
FuelFission:
Curium 245
Specific Power1 kg/MW
FuelFission:
Curium 245
ReactorZeta-Pinch
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorMagnetic Nozzle
Mini-Mag Orion (DRM-1)
Exhaust Velocity93,164 m/s
Specific Impulse9,497 s
Thrust642,000 N
Thrust Power29.9 GW
Mass Flow7 kg/s
Total Engine Mass119,046 kg
T/W0.55
Wet Mass731,924 kg
Dry Mass250,300 kg
Mass Ratio2.92 m/s
ΔV99,967 m/s
Specific Power4 kg/MW
Mini-Mag Orion (DRM-3)
Exhaust Velocity93,000 m/s
Specific Impulse9,480 s
Thrust642,000 N
Thrust Power29.9 GW
Mass Flow7 kg/s
Total Engine Mass199,600 kg
T/W0.33
Wet Mass788,686 kg
Dry Mass157,723 kg
Mass Ratio5.00 m/s
ΔV149,686 m/s
Specific Power7 kg/MW

The Mini-MagOrion is a sort of micro-fission Orion propulsion system. The idea was to make an Orion with weaker (and more reasonably sized) explosive pulses, using pulse charges that were not self contained (the full Orion pulse units were nothing less than nuclear bombs). Subcritical hollow spheres of curium-245 are compressed by a Z-pinch magnetic field until they explode. The sacrificial Z-pinch coil in each pulse charge is energized by an a huge external capacitor bank mounted in the spacecraft. So the pulse units are not bombs.

The explosion is caught by a superconducting magnetic nozzle.

More details are in the Realistic Designs section.

#### Zeta-Pinch Fusion

HOPE Z-Pinch
Propulsion SystemZ-Pinch Fusion
Exhaust Velocity189,780 m/s
Specific Impulse19,346 s
Thrust38,120 N
Thrust Power3.6 GW
Mass Flow0.20 kg/s
Total Engine Mass95,138 kg
T/W0.04
FuelDeuterium-Tritium fusion
+ Lithium6 fission
ReactorZeta-Pinch
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorMagnetic Nozzle
Wet Mass888,720 kg
Dry Mass552,000 kg
Mass Ratio1.61 m/s
ΔV90,380 m/s
Specific Power26 kg/MW
Firefly Starship
2013 design
ΔV2.698×107 m/s
(0.09c)
Wet Mass17,800 metric tons
Dry Mass2,365 metric tons
Mass Ratio7.526
PropulsionZ-Pinch DD Fusion
Exhaust Velocity1.289×107 m/s
Thrust1.9×106 N
Acceleration0.11 m/s
(0.01 g)
Accel time4 years
Coast time93 years
Decel time1 years

### PuFF Pulsed Fission Fusion

SpecificImpulse 20,000 sec 196,000 m/s 29,400 N 2.88 GW 96 kW/kg U-235 + D-T

The study authors were going to take a Hope Z-Pinch Fusion spacecraft and swap out its drive for the PuFF drive.

The idea is that while you can make some fuel undergo nuclear fission, and you can make other fuel undergo nuclear fusion, wouldn't it be nice to make some fuel do both? After all, a standard nuclear fusion warhead is a slug of fusion fuel that is ignited by the detonation of a small nuclear fission warhead.

Refer to the diagram at right.

The target is a charge of fission/fusion fuel, composed of Uranium-235 fission fuel and Deuterium-Tritium fusion fuel. The charge is held at the ignition point by some strong holder.

A ring of liquid lithium sprayers (Li Injectors) are aimed at the target. They spray a cone-shaped plume of liquid lithium (Li Shell) with the cone apex located at the target. Oh, did I mention that the sprayers are connected to the anode of the power system capacitor (LTDs) so they and the lithium shell are charged to two mega-volts? The target holder is connected to the cathode.

When the liquid lithium hits the target the circuit is closed, and the target is electrocuted by two mega-amps at two mega-volts (also totally draining the power system capacitor). This is 4 terawatts (4×1012 watts). Lorentz force (j×B) produced by the current and magnetic field savagely squeezes the fuel charge to one-tenth its original size. This makes the uranium achieve criticality.

Only some of the uranium undergoes nuclear fission like an atom bomb (which it is). This heats the D-T fuel hot enough to initiate nuclear fusion.

Neutrons from the fusion reaction ignites more of the uranium into a fission reaction. The heat from the fission boosts the fusion rate. Rinse-Lather-Repeat. This is called a Fission-Fusion Cascade. The fission to fusion cycle keeps cascading until all the fuel is burnt.

The energy from the cascade turns the liquid lithium into plasma. The plume of charged plasma from the cascade is ejected by the magnetic exhaust nozzle. In addition to creating thrust, the nozzle also harvests some of the exhaust energy to charge up the primary power system capacitors for the subsequent pulse.

Each fuel charge detonation takes several microseconds to cascade to full burnout. Detonations are repeated up to a rate of 100 Hz. The report notes that much analysis and experimentation is needed to find the optimum detonation frequency and fuel charge size.

The specific impulse and thrust can shift gears by modifying the amount of lithium injected.

Initially the power system capacitors are empty. For the first charge of the new burn an onboard SP-100 nuclear reactor laboriously charges them up. Subsequent capacitor recharges are by harvesting exhaust energy.

Left as an excercise for the reader is what the heck do you make the target holder out of so it is not obliterated by the fission and fusion explosions.

A - Target
Charge of fission/fusion fuel
B - Linear Transformer Drivers (LTD)
Pulsed power storage (capacitors), discharge, and compression system
C - Magnetic Nozzle (MN)
Directs fission/fusion products into exhaust for thrust. Recovers energy for next pulse.
D - Recharge System
Pulse generation and onboard power storage/generation
E - Lithium Injectors
Lithium tankage / distribution system to provide target liner (cone of liquid lithium) and power conduction path (when it touches the target)
F - Target Storage / Dispenser
Maintains targets in non-critical configuration (so the uranium doesn't explode prematurely), injects into nozzle

### Medusa

Medusa
Exhaust velocity490,000 m/s
to 980,000 m/s

Medusa is driven by the detonation of nuclear charges like Orion, except the charges are set off in front of the spacecraft instead of behind. The spacecraft trails behind a monstrously huge parachute shaped sail (about 500 meters). The sail intercepts the energy from the explosion. Medusa performs better than the classical Orion design because its pusher plate intercepts more of the bomb's blast, its shock-absorber stroke is much longer, and all its major structures are in tension and hence can be quite lightweight. It also scales down better. The nuclear charges will be from 0.025 kilotons to 2.5 kilotons.

The complicated stroke cycle is to smooth out the impulses from each blast, transforming it from a neck-braking jerk into a prolonged smooth acceleration.

Jondale Solem calculates that the specific impulse is a function of the mass and yield of the nuclear charges, while the thrust is a function of the yield and explosion repetition rate. In this case, the mass of the nuclear charge is the mass of "propellant".

Remarkably the mass of the spinnaker (sail) is independent of the size of its canopy or the number or length of its tethers. This means the canopy can be made very large (so the bomb blast radiation does not harm the canopy) and the tethers can be made very long (so the bomb blast radiation does not harm the crew). The mass of the spinnaker is directly proportional to the bomb yield and inversely proportional to the number of tethers.

### Inertial Confinement

IC-Fusion
Exhaust Velocity10,000,000 m/s
Specific Impulse1,019,368 s
Thrust100,000,000 N
Thrust Power500.0 TW
Mass Flow10 kg/s
Total Engine Mass1,000,000 kg
T/W10
FuelProton-Proton
Fusion
Specific Power2.00e-03 kg/MW

A pellet of fusion fuel is bombarded on all sides by strong pulses from laser or particle accelerators. Some use a two dimensional ring of lasers like a proverbial circular firing squad. Others expand it into a three dimensional spherical firing squad. The beams implode the pellet, raising the density and temperature to the point where a fusion reaction ignites.

The inertia of the fuel holds it together long enough for most of it to undergo fusion, instead of using a magnetic bottle as in Magnetic Confinement fusion.

The spherical arrangement of lasers would have a gap in it for the exhaust nozzle.

VISTA
Propulsion SystemIC Fusion
Exhaust Velocity170,000 m/s
Specific Impulse17,329 s
Thrust240,000 N
Thrust Power20.4 GW
Mass Flow1 kg/s
FuelDeuterium-Tritium
Fusion
ReactorInertial Confinement
Laser
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorMagnetic Nozzle
Wet Mass6,000,000 kg
Dry Mass1,835,000 kg
Mass Ratio3.27 m/s
Width170 m
Height100 m

#### Magneto Inertial Fusion

Magneto Inertial Fusion
Both
Exhaust Velocity50,420 m/s
Specific Impulse5,140 s
FuelDeuterium-Deuterium
Fusion
ReactorMagneto-Inertial
Confinement
RemassLithium
Remass AccelThermal Accel:
Reaction Heat
Low Gear
Thrust103 N
Thrust Power2.6 MW
Mass Flow2.00×10-03 kg/s
Delay between
Fusion Pulses
180 seconds
High Gear
Thrust13,800 N
Thrust Power0.3 GW
Mass Flow0.27 kg/s
Delay between
Fusion Pulses
14 seconds

There are two main approaches to utilizing nuclear fusion, magnetic confinement and inertial confinement. Magnetic confinement uses titanic magnetic fields, inertial confinement is how fusion bombs explode (a third way would be stars shining by gravitational confinement, but we don't know how to generate artificial gravitational fields).

Inertial confinement ignites the fusion fuel by imploding a solid pellet of fuel with a circular firing squad of lasers or particle beams. This raises the density and temperature high enough for fusion ignition. It confines the burning fusion fuel by sheer inertia. That is, it is hoping that the burning fuel simply does not have enough time to expand below fusion density before all the fuel is burnt.

Magnetic confinement ignites using a magnetic field to squeeze a cloud of fusion fuel plasma until it is hot and dense enough to ignite. It confines the burning fusion fuel with the same magnetic field. More like tries to confine, the blasted plasma keeps wiggling out of the cracks in the magnetic field before it is all burnt.

As propulsion systems, both have major drawbacks.

Problem 1

Magnetic confinement requires huge (read: massive) electromagnets. The technique also has the problem of plasma instabilities (read: fusion plasma has thousands of different ways to wiggle out of the magnetic cage) which so far have defied any solution. Meaning that every time fusion researchers have devised a new magnetic cage, the blasted plasma finds two new ways of wiggling out.

Inertial confinement works well in bombs, but trying to do it in a small controlled fashion (read: so the fusion reaction does not vaporize everything in a one kilometer radius) has also defied any solution. The compressing laser or particle beams have such low efficiencies that tons of excess power is required. Timing all the beams so they strike at the same instant is a challenge.

Problem 2

Both approaches have a problem with getting the fusion reaction energy to heat the propellant. Magnetic confinement tries to use the actual fusion plasma as propellant, resulting in a ridiculously small mass flow and thus a tiny thrust.

Problem 3

Also, there is nothing in between the fusion reaction and the chamber walls, leading to severe damage to the walls. The escaping radiation harms the crew as well.

Magneto Inertial Fusion

Dr. John Slough and his associates have come up with a new technique that sort of combines the two conventional approaches: magneto inertial fusion (MIF). You can find their published papers on the subject here

A blob of FRC (field reversed configuration) plasma is created and injected axially into the chamber.

Simultaneously injected into the chamber is a "liner". The liner is a foil ring composed of lithium, about 0.2 meters in radius. Each liner will have a mass of 0.28 kg (minimum) to 0.41 kg.

As the liner travels axially down the chamber, electromagnets crush it down into a solid cylinder (the crush speed is about 3 kilometers per second, the cylinder will have a radius of 5 centimeters). This is timed so that the plasma blob (plasmoid) is in the center of the cylinder. The liner compresses the plasmoid and ignites the fusion reaction.

The fusion reaction vaporizes the lithium liner. The ionized lithium (plus the burnt fusion fuel) exits through a magnetic nozzle, providing thrust.

In other words it both ignites and confines the fusion fuel with a collapsing wall of solid metal. The metal is being squeezed by an external magnetic field even as the fusion reaction is raging, which does a much better job of confinent than simple inertia or a rubbery magnetic field.

Liner compression is a heck of a lot more efficent than using huge magnetic fields or batteries of laser beams. Translation: it uses way less power and the equipment has a far smaller mass cost. Problem 1 solved.

The lithium is also the propellant. Since it is tightly wrapped around the reaction, it is very efficient at getting the fusion reaction energy to heat the propellant. Problem 2 solved.

The lithium stands in between the reaction and the chamber walls, protecting the walls. It also absorbs much of the radiation, protecting the crew. Problem 3 solved.

So magneto inertial fusion solves the fusion ignition problem, the fusion heating the propellant problem, and the reaction damaging the chamber problem which are endemic to magnetic and inertial confinement fusion. And the engine has a far lower mass.

Since this is an open-cycle system, the exhaust acts as the heat radiator, so the spacecraft can get by with only a tiny radiator. The energy to run the magnets can be supplied by solar cell arrays. Since the compression is so efficient, this will work with several types of fusion fuel: D-T, D-D, and D-3He. D-D is probably preferred, since tritium is radioactive with a short half-life, and 3He is rare.

Please note that if you replace the magnetic nozzle with a magnetohydrodynamic (MHD) generator, the propulsion system is transformed into an electrical power generator. This could be used for ground based fusion power generators.

Dr. Slough et al worked up two spacecraft for a Mars mission. The first was optimized to have a high payload mass fraction. The second was optimized to have the fastest transit time. Both were capable of a direct abort and return. The "Low Gear" engine is the study author's opinion of an engine easily achievable with current technology (that is, achievable fusion yields). The "High Gear" engine is a bit more speculative, but requiring only modest incremental improvements in technology.

Fusion Drive Rockets (FDR)
High Mass Fraction
EngineLow Gear
Transit Time90 days
Initial Mass90 mT
Specific Mass4.3 kg/kW
Shortest Transit Time
EngineHigh Gear
Transit Time30 days
Initial Mass153 mT
Specific Mass0.38 kg/kW

#### Plasma Jet Magneto Inertial

Plasma Jet Magnetio Inertial
ConMedOpt
Base Parameters
Mass of
plasma (g)
2.21.51.0
Efficiency of rail
& θ-pinch guns
0.3
Initial jet
velocity (km/s)
2007501500
Heat fraction
for 2nd power
0.010.001
Firing
Frequency (Hz)
200
Fusion
Gain
50
Target ΔV (c)0.08
Target Burn
Time (years)
4
Nozzle
Efficiency
0.84
Resulting Ship Parameters
Fuel Mass (t)55,50337,84325,228
Exhaust
Velocity (km/s)
1189.794461.278922.48
Specfic
Impulse (s)
121,284454,768909,529
Thrust (MN)0.521.341.78
Thrust/
Fuel Mass (N/kg)
0.03060.04490.0674
Jet Power (GW)311.432985.457961.07
Alpha (MW/kg)0.00560.07890.3156
Waste
Heat (GW)
15.14145.13387
Mass (t)
30029007732

This is basically the Magneto Inertial Fusion concept where the foil rings have been replaced by jets of plasma.

A blob of fusion fuel plasma is injected into the center of the reaction chamber. It is bombarded by a spherical firing squad much like classic inertial confinement fusion. The difference is:

1. The fusion fuel is a blob of plasma, not a solid pellet.
2. The fusion fuel plasma blob is magnetized.
3. The firing squad does not fire lasers or particle beams. Instead it fires cylindrical jets of plasma. The plasma is made from some element with a high atomic weight, so it has some serious momentum and inertia

In the table there are three columns for three estimates of the performance of an actual engine. These are labeled CON (Conservative), MED (Medium), and OPT (Optimistic). The report notes that the Medium column is probably good enough for an unmanned interstellar probe. The Conservative column is probably good enough for missions within the solar system.

• Mass Of Plasma: mass of the fusion fuel blob
• Efficiency of rail & θ-pinch guns: efficiency of the railguns shooting the plasma liner jets and the theta-pinch guns creating the fusion fuel blob
• Initial Jet Velocity: how fast the plasma liner jets are imploding
• Heat Fraction for 2nd Power: fraction of the total rejected heat that is being used for the secondary power needs of the spacecraft. Meaning that some of the waste heat will be sent through a generator to make power for the ship's avionics and whatnot
• Firing Frequency: how many fusion detonations are ignited per second
• Fusion Gain: how many times bigger is the fusion energy compared to the input energy. The return on your investment, in other words
• Target ΔV: the delta-V requirements the report assumes will be needed by the proposed interstellar mission
• Target Burn Time: the burn time requirements the report assumes will be needed by the proposed interstellar mission
• Nozzle Efficency: efficiency of the magnetic nozzle
• Fuel Mass: The total mass of fuel needed for the mission. This is also a first approximation of the ship's wet mass, since with such outrageous delta-V requirements the fuel mass will dominate the total mass
• Exhaust Velocity: what it says
• Specific Impulse: what it says
• Thrust: what it says
• Thrust/Fuel Mass: Thrust to total fuel mass ratio, which is pretty darn close to thrust to mass ratio
• Jet Power: what it says
• Alpha: power to mass ratio
• Waste Heat: amount of the power that turns up as waste heat and must be gotten rid of before the ship melts
• Radiator Mass: mass of the heat radiators required to cope with the waste heat

As the jets converge on the fuel at 750 kilometers per second they merge to form a spherical "liner". The liner collapses, squeezing the fusion fuel like a nutcracker from hell.

Meanwhile as the fusion fuel is squeezed, so is its magnetic field. The density of the magnetic field increases to a point where is makes a conventional magnetic-confinement fusion engine look anemic.

The shock where the imploding liner contacts the surface of the fuel blob heats it up. The liner also compresses the fuel blob, and soon fusion will be ignited. The internal magnetic field helps keep it confined long enough to burn all the fuel.

The exploding fusion blob hits the magnetic nozzle, compressing the nozzle's magnetic field. This acts like a trampoline, making the fusion plasma rebound out the exhaust nozzle, creating thrust. Which is the purpose of all rocket engines. Meanwhile some of the energy in the nozzle field compression can be harvested to charge up the capacitors for the next round.

The main advantage this propulsion system has over inertial or magnetic confinement is a drastically lower power requirement. Lasers, particle beam accelerators, or giant magnetics are power hogs. In this system the liner plasma jets can be lauched with relatively low powered rail-guns. This means you do not need tons and tons of capacitors to hold the huge jolts of electricity the other systems demand.

I will admit I am a little fuzzy on what advantage this system has over the metal foil type Magneto Inertial Fusion. I'm looking into it.

### Antimatter Bottle

This section has been moved here

### Antimatter catalyzed

Nuclear fission pulse drives like Orion scale up well, since it is relatively easy to design a bigger bomb than the last one. However, physics seem to prevent the creation of a nuclear device with a yield smaller than about 1/100 kiloton (10 tons, 42 GJ) and a fissionable material mass under 25 kilograms. This is due to critical mass restraints.

However, if a tiny sub-critical bit of fissionable material is bombarded by a few antiprotons, it will indeed create a tiny nuclear explosion. The antiprotons annihilate protons in uranium atoms, the energy release splits the atoms, creating a shower of neutrons, and a normal chain reaction ensues. Using antiprotons, yields smaller than 1/100 kiloton can be achieved. This can be used to create Antimatter catalyzed nuclear pulse propulsion

#### AIM

AIM
Exhaust Velocity598,000 m/s
Specific Impulse60,958 s
Thrust55 N
Thrust Power16.4 MW
Mass Flow1.00e-04 kg/s
FuelHelium3-Deuterium
Fusion
ReactorAntimatter Catalyzed
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorMagnetic Nozzle

Antiproton-initiated Microfusion. Inertial Confinement Fusion. See here.

#### ACMF

ICAN-II
Propulsion SystemACMF
Exhaust Velocity132,435 m/s
Specific Impulse13,500 s
Thrust180,000 N
Thrust Power11.9 GW
Mass Flow1 kg/s
Total Engine Mass27,000 kg
T/W0.68
FuelFission:
Uranium 235
ReactorAntimatter Catalyzed
RemassSilicon Carbide
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorAblative Nozzle
Wet Mass707,000 kg
Dry Mass345,000 kg
Mass Ratio2.05 m/s
ΔV95,020 m/s
Specific Power2 kg/MW

Antiproton-catalyzed microfission, inertial confinement fission. See here.

Fuel pellets have 3.0 grams of nuclear fuel (molar ratio of 9:1 of Deuterium:Uranium 235) coated with a spherical shell of 200 grams of lead. The lead shell is to convert the high energy radiation into a form more suited to be absorbed by the propellant. Each pellet produces 302 gigajoules of energy (about 72 tons of TNT) and are fired off at a rate of 1 Hz (one per second). The pellet explodes when it is struck by a beam containing about 1×1011 antiprotons.

A sector of a spherical shell of 4 meters radius is centered on the pellet detonation point. The shell is the solid propellant, silicon carbide (SiC), ablative propellant. The missing part of the shell constitutes the exhaust nozzle. Each fuel pellet detonation vaporizes 0.8 kilograms of propellant from the interior of the shell, which shoots out the exhaust port at 132,000 meters per second. This produces a thrust of 106,000 newtons.

The Penn State ICAN-II spacecraft was to have an ACMF engine, a delta-V capacity of 100,000 m/s, and a dry mass of 345 metric tons. The delta-V and exhaust velocity implied a mass ratio of 2.05. The dry mass and the mass ratio implied that the silicon carbide propellant shell has a mass of 362 metric tons. The wet mass and the thrust implied an acceleration of 0.15 m/s2 or about 0.015g. It can boost to a velocity of 25 km/sec in about three days. At 0.8 kilograms propellant ablated per fuel pellet, it would require about 453,000 pellets to ablat the entire propellant shell.

It carries 65 nanograms of antiprotons in the storage ring. At about 7×1014 antiprotons per nanogram, and 1×1011 antiprotons needed to ignite one fuel pellet, that's enough to ignite about 453,000 fuel pellets.

The system is very similar to Positron Ablative.

## Sail

Sail propulsion does not carry onboard reaction mass or does not use reaction mass. They are powered by a remote source, either the Sun or a satellite installation with a huge power supply and an equally huge laser/plasma beam.

Sail propulsion systems are Propellant-less Rockets, and thus is not subject to The Tyranny of the Rocket Equation.

### Electric Sail

This is a Propellant-less Rocket, and thus is not subject to The Tyranny of the Rocket Equation.

### Magnetic Sail

Magnetic Sail
Thrust per sail area0.001 N/km2
Thrust by Sol dist1/R2

This is a Propellant-less Rocket, and thus is not subject to The Tyranny of the Rocket Equation.

### M2P2

M2P2
Thrust per sail area0.001 N/km2
Thrust by Sol distConstant
Disk Inflates
as 1/R2
Plasma use0.25 kg/Day per N Thrust
Isp = 35,000

A Mini-magnetospheric plasma sail (M2P2) is a MagSail inflated by an injection of plasma, powered by the solar wind.

This is a Propellant-less Rocket, and thus is not subject to The Tyranny of the Rocket Equation.

### MagBeam

A MagBeam is Mini-magnetospheric plasma sail beam-powered by a remote helicon plasma beam installation.

This is a Propellant-less Rocket, and thus is not subject to The Tyranny of the Rocket Equation.

Report here. Alternatively the spacecraft can use a plasma magnet instead of a M2P2 to intercept the beam. With the current design, the spacecraft mass cannot be larger than about 10,000 kg (10 metric tons).

The installation is called a High Power Platform (HPP). The HPP does not have much range, so the spacecraft will require a second HPP at the destination in order to slow down. For a Mars mission the HPP fires for about four hours before the spacecraft is out of range. By that time the spacecraft is travelling at about 20,000 m/s, which is fast enough to get to Mars in 50 days flat. The range is about 1×107 meters (ten thousand kilometers).

After boosting a spacecraft, the HPP rotates the MagBeam in the opposite direction and uses it as an ion drive to move back into position. Newton's laws still hold, the recoil from the MagBeam is going to push the HPP way off base.

And I'm quite sure that at short ranges the MagBeam can be used as a weapon. Please note that when I say "short range", I mean "less than 50 meters or so."

It would also be a nifity thing for a warship to mount, so it can use it to boost missiles to ferocious velocities.

The main advantages seem to be increased acceleration levels on the spacecraft, and that one HPP propulsion unit can service multiple spacecraft. There are certain maneuvers that are impossible for low acceleration spacecraft, such as sub-orbital to orbital transfers, LEO to GEO transfers, LEO to escape velocity, and fast planetary missions.

Plasma beams as a general rule have short ranges. However, the system can take advantage of the fact that both the HPP and the spacecraft have magnetic fields. The MagBeam uses magnetic fields to focus the beam and the spacecraft has a MagSail to catch the beam. If they start off close enough to each other, the two magnetic field merge ("magnetic reconnection"), and gradually stretch as the spacecraft moves. This creates a long magnetic tunnel to confine the plasma stream, making the stream self-focusing.

This will be a problem when the HPP is faced with the task of slowing down an incoming spacecraft, since initially there will be no magnetic link. The spacecraft will have to temporarily inflate its MagSail, which can be done because it is an M2P2. Once the magnetic connection is made the M2P2 can be deflated to normal size.

Plasma will probably be argon or nitrogen. The beam range will a few thousand kilometers if the HPP or the beam passes through the ionosphere, tens of thousands of kilometers if in the magnetosphere. This is because of the ambient plasma and magnetic fields in the ionosphere.

Since the spacecraft does not carry its propellant, the standard rocket equation does not apply. Instead:

HPPe = (0.25 * M * deltaV * Ve ) / HPPeff

where:

• HPPe = electrical energy expended by HPP (joules)
• M = mass of spacecraft (kg)
• deltaV = delta V applied to spacecraft (m/s)
• HPPeff = efficiency of HPP at converting electricity into plasma energy (100% = 1.0, currently 0.6)

Mpb = HPPe / (0.5 * Ve2)

where:

• Mpb = mass of propellant expended in HPP beam (kg)
• HPPe = electrical energy expended by HPP (joules)
• Ve = velocity of HPP beam (m/s)

HPPpower = HPPe / Taccel

where:

• HPPpower = miminum power level of HPP power plant (watts)
• HPPe = electrical energy expended by HPP (joules)
• Taccel = duration of HPP beam usage (sec)

So if a HPP had to boost a 10,000 kg (10 metric ton) spacecraft to a deltaV of 3,000 m/s (3 km/s) using a plasma beam with a velocity of 19,600 m/s (2000 s) had only 300 seconds (5 minutes) to do so and had an efficiency of 0.6 (60%), then the electrical power used would be 2.5×1010 joules, the power plant would need a level of 82,000,000 watts (82 megawatts), and 127 kilograms of propellant would be expended.

### Photon Sail

Photon Sail
Thrust per sail area9 N/km2
Thrust by Sol dist1/R2

A Photon Sail is a sail powered by solar photons. Commonly called a "solar sail", but that common term does not make it clear if the sail is powered by solar photons, solar magnetic field, or solar wind.

This is a Propellant-less Rocket, and thus is not subject to The Tyranny of the Rocket Equation.

### Plasma Magnet

An plasma Magnet is a type of E-sail powered by solar wind.

This is a Propellant-less Rocket, and thus is not subject to The Tyranny of the Rocket Equation.

## Other

### ( Beer )

Beer
Thrust Power8 × 10-8 GW
Exhaust velocity83 m/s
Thrust84 n
T/W >1.0no

In The Makeshift Rocket (also known as A Bicycle Built for Brew), the old geezer cobbles together a crude rocket out of hogs-heads of pressurized beer in order to escape to an adjacent asteroid.

### Mass Driver

Mass Driver
Exhaust Velocity30,000 m/s
Specific Impulse3,058 s
Thrust20,000 N
Thrust Power0.3 GW
Mass Flow0.67 kg/s
Total Engine Mass150,000 kg
T/W0.01
Thermal eff.90%
Total eff.90%
Fuel800MWe input
RemassRegolith
Remass AccelElectromagnetic
Acceleration
Specific Power500 kg/MW

Mass drivers use electromagnetic accelerators to hurl mass. Much like an ion drive the "fuel" is electricity and the propellant is convenient matter. Better: ion drives want propellant that can be easily ionized, mass drivers don't care what you use for propellant.

There are actually two types: Integral Mass Drivers and External Mass Drivers.

INTEGRAL MASS DRIVERS: the electromagnetic accelerator is mounted on the spacecraft. Magnetic buckets filled with propellant, which is rock dust or anything else you can stuff into the bucket. The electromagnetic accelerator propels the bucket at high speed. At the end of the accelerator, the bucket is braked to a halt, but the propellant keeps flying. The propellant exits the accelerator and creates thrust on the spacecraft like any other rocket.

Integral mass drivers are popular with asteroid miners who want to nudge their claimed asteroid into more convenient orbits, since the rocks on the asteroid provide all the propellant you need for free. However, such asteroid moving operations may prompt the creation of a Spaceguard.

EXTERNAL MASS DRIVERS: the electromagnetic accelerator is mounted at a spaceport. The "propellant" is the spacecraft. The spacecraft is placed in a separate magnetic bucket or has hunks of ferrous metal incorporated into the ship's thrust frame. The accelerator throws the ship on its planned trajectory without the ship having to burn any fuel or reaction mass. The spaceport requires a large power source to energize the accelerator, and lots of bracing to dissipate the accelerator recoil.

External Mass Driver are sometimes called "electromagnetic catapults"

This is a Propellant-less Rocket, and thus is not subject to The Tyranny of the Rocket Equation.

In Gerard O'Neill's plan for L5 colonies, external mass drivers were located at lunar mining sites producing the raw materials for the colony. Instead of throwing spacecraft, they threw cannisters of raw materials (with no rocket engines at all). These were intercepted at the L5 point by a "catcher". So instead of needing a fleet of cargo rockets, you just needed a mass driver launcher, a catcher and lots of ferrous cannisters (which can be manufactured at the mining site out of local materials). The concept is called an inert cargo vessel.

A mass driver is an electromagnetic mass accelerator that is optimized for propulsion. If you optimize it as a weapon instead, you have a coil-gun or rail gun. In this case the "propellant" is a bullet or a cannon shell intended to perforate a hostile spacecraft. The weapons still have recoil and can be used as a crude propulsion system. You can do this with an external mass driver as well, turning a spaceport into a planetary fortress. One of the first SF authors to point this out was Robert Heinlein in The Moon is a Harsh Mistress.

### Photon

Photon
Exhaust Velocity299,792,458 m/s
Specific Impulse30,559,884 s
Fuel1.1TWe input
RemassPhotons
Remass AccelElectromagnetic
Acceleration

The exhaust is not a stream of matter. Instead it is a beam of Electromagnetic radiation, basically a large laser. The advantage is that it has the maximum possible exhaust velocity and thus the highest specific impulse. The disadvantage is the ludicrously high power requirements.

The momentum of a photon is p = E/c, where E is the energy of the photon. So the thrust delivered by a stream of photons is ∂p/∂t = ∂E/∂t/c. This boils down to:

F = P/c

P = F * c

where:

• F = thrust in Newtons
• P = power in watts
• c = speed of light in a vacuum (3e8 m/s)

In other words, one lousy Newton of thrust takes three hundred freaking megawatts!!

### Tachyon

This is pretty close to fringe physics. I know when you see the word "tachyon" you think "faster than light starship" but that is not what Dr. Cramer is speculating about here.

What it boils down to is an incredibly efficient propellant-less rocket. Mass ratios are worthless because the propellant mass is zero, this drive sneers at the Tyranny of the Rocket Equation.

Like a photon drive, it carries no propellant, it manufactures it out of electricity, as needed. The difference is:

1. the propellant is composed of tachyons, instead of photons as in the photon drive
2. it probably can create one newton of thrust with much less energy than three hundred megawatts

The problem is this drive runs afoul of Burnside's Advice. I know the tachyon drive is not reactionless, but it shares the same problem: it will give you Dirt Cheap Planet Crackers. You might be able to put a band-aid on the problem by dialing up the required energy per newton of thrust. But I fear the range of economically viable propulsion is very similar to the range of dirt cheap planet crackers.

## Watch the Heat

This section has been moved here

## Atomic Rockets notices

This week's featured addition is Spacecraft from RACE TO MARS

This week's featured addition is Space Rescue Vehicle

This week's featured addition is Dusty Plasma Fission Fragment Rocket